Overall engine efficiency rating for turbomachine engines

ABSTRACT

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and an overall engine efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include a low-pressure turbine. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine of the core engine and comprises a first rotational speed, the output of the gearbox is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-4.0. The overall engine efficiency rating is greater than or equal to 0.35GR1.5 and less than or equal to 0.7GR1.5.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/929,542, filed Sep. 2, 2022, which is a continuation of U.S. patentapplication Ser. No. 17/694,444, filed Mar. 14, 2022, now U.S. Pat. No.11,466,624, which claims the benefit of Italian Application No.102022000001613, filed Jan. 31, 2022. The prior applications areincorporated by reference herein.

ACKNOWLEDGMENT OF GOVERNMENT SUPPORT

The project leading to this application has received funding from theClean Sky 2 Joint Undertaking (JU) under grant agreement No. 945541. TheJU receives support from the European Union's Horizon 2020 research andinnovation programme and the Clean Sky 2 JU members other than theUnion.

FIELD

This disclosure relates generally to turbomachines comprising a gearboxand particularly to geared turbofan engines.

BACKGROUND

A turbofan engine includes a core engine that drives a bypass fan. Thebypass fan generates the majority of the thrust of the turbofan engine.The generated thrust can be used to move a payload (e.g., an aircraft).

In some instances, a turbofan engine is configured as a direct driveengine. Direct drive engines are configured such that a power turbine(e.g., a low-pressure turbine) of the core engine is directly coupled tothe bypass fan. As such, the power turbine and the bypass fan rotate atthe same rotational speed (i.e., the same rpm).

In other instances, a turbofan engine can be configured as a gearedengine. Geared engines include a gearbox disposed between andinterconnecting the bypass fan and power turbine of the core engine. Thegearbox, for example, allows the power turbine of the core engine torotate at a different speed than the bypass fan. Thus, the gearbox can,for example, allow the power turbine of the core engine and the bypassfan to operate at their respective rotational speeds for maximumefficiency and/or power production.

Despite certain advantages, geared turbofan engines can have one or moredrawbacks. For example, including a gearbox in a turbofan engineintroduces additional complexity to the engine. This can, for example,make engine development and/or manufacturing significantly moredifficult. As such, there is a need for improved geared turbofanengines. There is also a need for devices and methods that can be usedto develop and manufacture geared turbofan engines more efficientlyand/or precisely.

BRIEF DESCRIPTION

Aspects and advantages of the disclosed technology will be set forth inpart in the following description, or may be obvious from thedescription, or may be learned through practice of the technologydisclosed in the description.

Various turbomachine engines and gear assemblies are disclosed herein.The disclosed turbomachine engines comprise a gearbox. The disclosedturbomachine engines are characterized or defined by an overall engineefficiency rating. The overall engine efficiency rating equals

${{Q\left( \frac{D^{1.56}}{\tau} \right)}^{1.53}N^{2}},$

where Q is a gearbox oil flow rate an inlet of the gearbox measured ingallons per minute at a max takeoff condition, D is a diameter of thefan blades measured in inches, T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition, and N is anumber of rotating blade stages of a low-pressure turbine. Values forthe overall engine efficiency rating identify key engine requirementsaffecting the overall architecture. An engine architecture based, atleast in part, on this value, can enable early optimization of majorengine components, thereby benefiting the overall architecture. Thus,the overall engine efficiency rating may also be used, for example, toaid the development of the fan, gearbox, and/or low-pressure turbine inrelation to other engine parameters. The overall engine efficiencyrating thus provides improved turbomachine engines and/or can helpsimplify one or more complexities of geared turbomachine enginedevelopment.

In one example, a turbomachine engine includes: a fan assembly includinga plurality of fan blades; a vane assembly including a plurality ofvanes disposed aft of the plurality of fan blades; a core engineincluding a low-pressure turbine; a gearbox including an input and anoutput, wherein the input of the gearbox is coupled to the low-pressureturbine of the core engine and comprises a first rotational speed,wherein the output of the gearbox is coupled to the fan assembly and hasa second rotational speed, and wherein a gear ratio of the firstrotational speed to the second rotational speed is within a range of3.2-4.0; and an overall engine efficiency rating of 0.57-8.0, whereinthe overall engine efficiency rating equals

${{Q\left( \frac{D^{1.56}}{\tau} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

In another example, a turbomachine engine includes: a fan casing; a fanassembly disposed within the fan case and including a plurality of fanblades; a vane assembly disposed within the fan case and including aplurality of vanes disposed aft of the plurality of fan blades; a coreengine including a low-pressure compressor, a high-pressure compressor,a combustor, a high-pressure turbine, and a low-pressure turbine,wherein the high-pressure turbine is coupled to the high-pressurecompressor via a high-speed shaft, and wherein the low-pressure turbineis coupled to the low-speed compressor via a low-speed shaft; a gearboxincluding an input and an output, wherein the input of the gearbox iscoupled to the low-speed shaft and comprises a first rotational speed,wherein the output of the gearbox is coupled to the fan assembly and hasa second rotational speed, and wherein a gear ratio of the firstrotational speed to the second rotational speed is within a range of3.25-3.75; and an overall engine efficiency rating of 0.59-7.3, whereinthe overall engine efficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotor stages of the low-pressureturbine.

In another example, a turbomachine engine includes: a ducted fanassembly including a plurality of fan blades; a ducted vane assemblyincluding a plurality of vanes, wherein the plurality of vanes isconfigured to receive a first portion of airflow from the plurality offan blades; a core engine configured to receive a second portion of theairflow from the plurality of fan blades, wherein the core engineincludes a low-pressure compressor, a high-pressure compressor, acombustor, a high-pressure turbine, and a low-pressure turbine; agearbox comprising a gear ratio within a range of 3.2-4.0; and anoverall engine efficiency rating of 0.8-3.0.

In another example, a turbomachine engine includes: a nacelle; a fanassembly disposed within the nacelle and including a plurality of fanblades arranged in a single blade row; a vane assembly disposed withinthe nacelle and including a plurality of vanes arranged in a single vanerow and disposed aft of the plurality of fan blades; a core engineincluding a first compressor section, a second compressor section, afirst turbine section, and a second turbine section; a first shaftcoupling the first turbine section to the first compressor section; asecond shaft coupling the second turbine section to the secondcompressor section; a gearbox including an input and an output, whereinthe input of the gearbox is coupled to the first shaft and comprises afirst rotational speed, wherein the output of the gearbox is coupled tothe fan assembly and has a second rotational speed, which is less thanthe first rotational speed, and wherein a gear ratio of the gearbox iswithin a range of 3.2-4.0; and an overall engine efficiency rating of0.57-8.0, wherein the gearbox efficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thefirst turbine section and equals 4.

In another example, a turbomachine engine includes: a fan assemblyincluding a plurality of fan blades; a vane assembly including aplurality of vanes disposed aft of the plurality of fan blades; a coreengine including a low-pressure turbine; a gearbox including an inputand an output, wherein the input of the gearbox is coupled to thelow-pressure turbine of the core engine and comprises a first rotationalspeed, wherein the output of the gearbox is coupled to the fan assemblyand has a second rotational speed, and wherein a gear ratio (GR) of thefirst rotational speed to the second rotational speed is within a rangeof 3.2-4.0; and an overall engine efficiency rating greater than0.1GR^(1.5) and less than GR^(1.5), wherein the overall engineefficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

In another example, a turbomachine engine includes: a fan assemblyincluding a plurality of fan blades; a vane assembly including aplurality of vanes disposed aft of the plurality of fan blades; a coreengine including a low-pressure turbine; a gearbox including an inputand an output, wherein the input of the gearbox is coupled to thelow-pressure turbine of the core engine and comprises a first rotationalspeed, wherein the output of the gearbox is coupled to the fan assemblyand has a second rotational speed, and wherein a gear ratio (GR) of thefirst rotational speed to the second rotational speed is within a rangeof 2.0-2.9; and an overall engine efficiency rating greater than0.1GR^(1.5) and less than GR^(1.5), wherein the overall engineefficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

In another example, a turbomachine engine includes: a fan assemblyincluding a plurality of fan blades; a vane assembly including aplurality of vanes disposed aft of the plurality of fan blades; a coreengine including a low-pressure turbine; a gearbox including an inputand an output, wherein the input of the gearbox is coupled to thelow-pressure turbine of the core engine and comprises a first rotationalspeed, wherein the output of the gearbox is coupled to the fan assemblyand has a second rotational speed, and wherein a gear ratio (GR) of thefirst rotational speed to the second rotational speed is within a rangeof 2.0-4.0; and an overall engine efficiency rating greater than0.35GR^(1.5) and less than 0.7GR^(1.5), wherein the overall engineefficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

These and other features, aspects, and/or advantages of the presentdisclosure will become better understood with reference to the followingdescription and the claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateexamples of the disclosed technology and, together with the description,serve to explain the principles of the disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a cross-sectional schematic illustration of an example of aturbomachine engine configured with an open rotor propulsion system.

FIG. 2 is a cross-sectional schematic illustration of an example of aturbomachine engine configured with an open rotor propulsion system.

FIG. 3 is a cross-sectional schematic illustration of an example of aturbomachine engine configured with a ducted propulsion system.

FIG. 4 is a cross-sectional schematic illustration of an example of aturbomachine engine configured with a ducted propulsion system.

FIG. 5 is a cross-sectional schematic illustration of an example of acounter-rotating low-pressure turbine of a turbomachine engine, thelow-pressure turbine having a 3×3 configuration.

FIG. 6 is a cross-sectional schematic illustration of an example of acounter-rotating low-pressure turbine of a turbomachine engine, thelow-pressure turbine having a 4×3 configuration.

FIG. 7A is a graph depicting an exemplary range of gearbox efficiencyratings relative to an exemplary range of gear ratios for a turbomachineengine.

FIG. 7B is a graph depicting another exemplary range of gearboxefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 7C is a graph depicting another exemplary range of gearboxefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 8 is a chart depicting various engine parameters of severalexemplary turbomachine engines, including a gearbox efficiency rating.

FIG. 9 is a cross-sectional schematic illustration of an example of agearbox configuration for a turbomachine engine.

FIG. 10 is a cross-sectional schematic illustration of an example of agearbox configuration for a turbomachine engine.

FIG. 11 is a cross-sectional schematic illustration of an example of agearbox configuration for a turbomachine engine.

FIG. 12 is a cross-sectional schematic illustration of an example of agearbox configuration for a turbomachine engine.

FIG. 13 is a cross-sectional schematic illustration of an example of agearbox configuration for a turbomachine engine.

FIG. 14 is a schematic diagram of an exemplary lubricant systemsupplying lubricant to an engine component.

FIG. 15 is a schematic diagram of the lubricant system configured tosupply lubricant to a gearbox.

FIG. 16A is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 16B is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 16C is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 16D is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 17A is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 17B is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 17C is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 18A is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 18B is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 19A is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 19B is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 19C is a graph depicting an exemplary range of overall engineefficiency ratings relative to an exemplary range of gear ratios for aturbomachine engine.

FIG. 20 is a chart depicting various engine parameters of severalexemplary turbomachine engines, including an overall engine efficiencyrating.

DETAILED DESCRIPTION

Reference now will be made in detail to examples of the disclosedtechnology, one or more examples of which are illustrated in thedrawings. Each example is provided by way of explanation of thedisclosed technology, not a limitation of the disclosure. In fact, itwill be apparent to those skilled in the art that various modificationsand variations can be made in the present disclosure without departingfrom the scope or spirit of the disclosure. For instance, featuresillustrated or described as part of one example can be used with anotherexample to yield a still further example. Thus, it is intended that thepresent disclosure covers such modifications and variations as comewithin the scope of the appended claims and their equivalents.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations.

As used herein, the terms “first,” “second,” and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify the location or importance of the individualcomponents.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a,” “an,” and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about,” “approximately,” and “substantially,” is not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,5, 10, 15, or 20 percent margin in either individual values, range(s) ofvalues and/or endpoints defining range(s) of values.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

One or more components of the turbomachine engine or gear assemblydescribed herein below may be manufactured or formed using any suitableprocess, such as an additive manufacturing process, such as a 3-Dprinting process. The use of such a process may allow such components tobe formed integrally, as a single monolithic component, or as anysuitable number of sub-components. In particular, the additivemanufacturing process may allow such components to be integrally formedand include a variety of features not possible when using priormanufacturing methods. For example, the additive manufacturing methodsdescribed herein enable the manufacture of heat exchangers having uniquefeatures, configurations, thicknesses, materials, densities, fluidpassageways, headers, and mounting structures that may not have beenpossible or practical using prior manufacturing methods. Some of thesefeatures are described herein.

Referring now to the drawings, FIG. 1 is an example of an engine 100including a gear assembly 102 according to aspects of the presentdisclosure. The engine 100 includes a fan assembly 104 driven by a coreengine 106. In various examples, the core engine 106 is a Brayton cyclesystem configured to drive the fan assembly 104. The core engine 106 isshrouded, at least in part, by an outer casing 114. The fan assembly 104includes a plurality of fan blades 108. A vane assembly 110 extends fromthe outer casing 114 in a cantilevered manner. Thus, the vane assembly110 can also be referred to as an unducted vane assembly. The vaneassembly 110, including a plurality of vanes 112, is positioned inoperable arrangement with the fan blades 108 to provide thrust, controlthrust vector, abate or re-direct undesired acoustic noise, and/orotherwise desirably alter a flow of air relative to the fan blades 108.

In some examples, the fan assembly 104 includes eight (8) to twenty (20)fan blades 108. In particular examples, the fan assembly 104 includesten (10) to eighteen (18) fan blades 108. In certain examples, the fanassembly 104 includes twelve (12) to sixteen (16) fan blades 108. Insome examples, the vane assembly 110 includes three (3) to thirty (30)vanes 112. In certain examples, the vane assembly 110 includes an equalor fewer quantity of vanes 112 to fan blades 108. For example, inparticular examples, the engine 100 includes twelve (12) fan blades 108and ten (10) vanes 112. In other examples, the vane assembly 110includes a greater quantity of vanes 112 to fan blades 108. For example,in particular implementations, the engine 100 includes ten (10) fanblades 108 and twenty-three (23) vanes 112.

In certain examples, such as depicted in FIG. 1 , the vane assembly 110is positioned downstream or aft of the fan assembly 104. However, itshould be appreciated that in some examples, the vane assembly 110 maybe positioned upstream or forward of the fan assembly 104. In stillvarious examples, the engine 100 may include a first vane assemblypositioned forward of the fan assembly 104 and a second vane assemblypositioned aft of the fan assembly 104. The fan assembly 104 may beconfigured to desirably adjust pitch at one or more fan blades 108, suchas to control thrust vector, abate or re-direct noise, and/or alterthrust output. The vane assembly 110 may be configured to desirablyadjust pitch at one or more vanes 112, such as to control thrust vector,abate or re-direct noise, and/or alter thrust output. Pitch controlmechanisms at one or both of the fan assembly 104 or the vane assembly110 may co-operate to produce one or more desired effects describedabove.

In certain examples, such as depicted in FIG. 1 , the engine 100 is anun-ducted thrust producing system, such that the plurality of fan blades108 is unshrouded by a nacelle or fan casing. As such, in variousexamples, the engine 100 may be configured as an unshrouded turbofanengine, an open rotor engine, or a propfan engine. In particularexamples, the engine 100 is an unducted rotor engine with a single rowof fan blades 108. The fan blades 108 can have a large diameter, such asmay be suitable for high bypass ratios, high cruise speeds (e.g.,comparable to aircraft with turbofan engines, or generally higher cruisespeed than aircraft with turboprop engines), high cruise altitude (e.g.,comparable to aircraft with turbofan engines, or generally higher cruisespeed than aircraft with turboprop engines), and/or relatively lowrotational speeds.

The fan blades 108 comprise a diameter (D_(fan)). It should be notedthat for purposes of illustration only half of the D_(fan) is shown(i.e., the radius of the fan). In some examples, the D_(fan) is 72-216inches. In particular examples the D_(fan) is 100-200 inches. In certainexamples, the D_(fan) is 120-190 inches. In other examples, the D_(fan)is 72-120 inches. In yet other examples, the D_(fan) is 50-80 inches.

In some examples, the fan blade tip speed at a cruise flight conditioncan be 650 to 900 fps, or 700 to 800 fps. A fan pressure ratio (FPR) forthe fan assembly 104 can be 1.04 to 1.10, or in some examples 1.05 to1.08, as measured across the fan blades at a cruise flight condition.

Cruise altitude is generally an altitude at which an aircraft levelsafter climb and prior to descending to an approach flight phase. Invarious examples, the engine is applied to a vehicle with a cruisealtitude up to approximately 65,000 ft. In certain examples, cruisealtitude is between approximately 28,000 ft. and approximately 45,000ft. In still certain examples, cruise altitude is expressed in flightlevels (FL) based on standard air pressure at sea level, in which acruise flight condition is between FL280 and FL650. In another example,cruise flight condition is between FL280 and FL450. In still certainexamples, cruise altitude is defined based at least on barometricpressure, in which cruise altitude is between approximately 4.85 psiaand approximately 0.82 psia based on a sea-level pressure ofapproximately 14.70 psia and sea-level temperature at approximately 59degrees Fahrenheit. In another example, cruise altitude is betweenapproximately 4.85 psia and approximately 2.14 psia. It should beappreciated that in certain examples, the ranges of cruise altitudedefined by pressure may be adjusted based on a different referencesea-level pressure and/or sea-level temperature.

The core engine 106 is generally encased in outer casing 114 definingone-half of a core diameter (D_(core)), which may be thought of as themaximum extent from the centerline axis (datum for R). In certainexamples, the engine 100 includes a length (L) from a longitudinally (oraxial) forward end 116 to a longitudinally aft end 118. In variousexamples, the engine 100 defines a ratio of L/D_(core) that provides forreduced installed drag. In one example, L/D_(core) is at least 2. Inanother example, L/D_(core) is at least 2.5. In some examples, theL/D_(core) is less than 5, less than 4, and less than 3. In variousexamples, it should be appreciated that the L/D_(core) is for a singleunducted rotor engine.

The reduced installed drag may further provide for improved efficiency,such as improved specific fuel consumption. Additionally, oralternatively, the reduced drag may provide for cruise altitude engineand aircraft operation at or above Mach 0.5. In certain examples, theL/D_(core), the fan assembly 104, and/or the vane assembly 110separately or together configure, at least in part, the engine 100 tooperate at a maximum cruise altitude operating speed betweenapproximately Mach 0.55 and approximately Mach 0.85; or betweenapproximately 0.72 to 0.85 or between approximately 0.75 to 0.85.

Referring still to FIG. 1 , the core engine 106 extends in a radialdirection (R) relative to an engine centerline axis 120. The gearassembly 102 receives power or torque from the core engine 106 through apower input source 122 and provides power or torque to drive the fanassembly 104, in a circumferential direction C about the enginecenterline axis 120, through a power output source 124.

The gear assembly 102 of the engine 100 can include a plurality ofgears, including an input and an output. The gear assembly can alsoinclude one or more intermediate gears disposed between and/orinterconnecting the input and the output. The input can be coupled to aturbine section of the core engine 106 and can comprise a firstrotational speed. The output can be coupled to the fan assembly and canhave a second rotational speed. In some examples, a gear ratio of thefirst rotational speed to the second rotational speed is less than orequal to four (e.g., within a range of 2.0-4.0). In other examples, agear ratio of the first rotational speed to the second rotational speedis greater than four (e.g., within a range of 4.1-14.0).

The gear assembly 102 (which can also be referred to as “a gearbox”) cancomprise various types and/or configurations. For example, in someinstances, the gearbox is an epicyclic gearbox configured in a star gearconfiguration. Star gear configurations comprise a sun gear, a pluralityof star gears (which can also be referred to as “planet gears”), and aring gear. The sun gear is the input and is coupled to the power turbine(e.g., the low-pressure turbine) such that the sun gear and the powerturbine rotate at the same rotational speed. The star gears are disposedbetween and interconnect the sun gear and the ring gear. The star gearsare rotatably coupled to a fixed carrier. As such, the star gears canrotate about their respective axes but cannot collectively orbitrelative to the sun gear or the ring gear. As another example, thegearbox is an epicyclic gearbox configured in a planet gearconfiguration. Planet gear configurations comprise a sun gear, aplurality of planet gears, and a ring gear. The sun gear is the inputand is coupled to the power turbine. The planet gears are disposedbetween and interconnect the sun gear and the ring gear. The planetgears are rotatably coupled to a rotatable carrier. As such, the planetgears can rotate about their respective axes and also collectivelyrotate together with the carrier relative to the sun gear and the ringgear. The carrier is the output and is coupled to the fan assembly. Thering gear is fixed from rotation.

In some examples, the gearbox is a single-stage gearbox (e.g., FIGS.10-11 ). In other examples, the gearbox is a multi-stage gearbox (e.g.,FIGS. 9 and 12 ). In some examples, the gearbox is an epicyclic gearbox.In some examples, the gearbox is a non-epicyclic gearbox (e.g., acompound gearbox—FIG. 13 ).

As noted above, the gear assembly can be used to reduce the rotationalspeed of the output relative to the input. In some examples, a gearratio of the input rotational speed to the output rotational speed iswithin a range of 2-4. For example, the gear ratio can be 2-2.9, 3.2-4,or 3.25-3.75). In some examples, a gear ratio of the input rotationalspeed to the output rotational speed is greater than 4.1. For example,in particular instances, the gear ratio is within a range of 4.1-14.0,within a range of 4.5-14.0, or within a range of 6.0-14.0. In certainexamples, the gear ratio is within a range of 4.5-12 or within a rangeof 6.0-11.0. As such, in some examples, the fan assembly can beconfigured to rotate at a rotational speed of 700-1500 rpm at a cruiseflight condition, while the power turbine (e.g., the low-pressureturbine) is configured to rotate at a rotational speed of 2,500-15,000rpm at a cruise flight condition. In particular examples, the fanassembly can be configured to rotate at a rotational speed of 850-1350rpm at a cruise flight condition, while the power turbine is configuredto rotate at a rotational speed of 5,000-10,000 rpm at a cruise flightcondition.

Various gear assembly configurations are depicted schematically in FIGS.9-13 . These gearboxes can be used with any of the engines disclosedherein, including the engine 100. Additional details regarding thegearboxes are provided below.

FIG. 2 shows a cross-sectional view of an engine 200, which isconfigured as an example of an open rotor propulsion engine. The engine200 is generally similar to the engine 100 and corresponding componentshave been numbered similarly. For example, the gear assembly of theengine 100 is numbered “102” and the gear assembly of the engine 200 isnumbered “202,” and so forth. In addition to the gear assembly 202, theengine 200 comprises a fan assembly 204 that includes a plurality of fanblades 208 distributed around the engine centerline axis 220. Fan blades208 are circumferentially arranged in an equally spaced relation aroundthe engine centerline axis 220, and each fan blade 208 has a root 225and a tip 226, and an axial span defined therebetween, as well as acentral blade axis 228.

The core engine 206 includes a compressor section 230, a combustionsection 232, and a turbine section 234 (which may be referred to as “anexpansion section”) together in a serial flow arrangement. The coreengine 206 extends circumferentially relative to an engine centerlineaxis 220. The core engine 206 includes a high-speed spool that includesa high-speed compressor 236 and a high-speed turbine 238 operablyrotatably coupled together by a high-speed shaft 240. The combustionsection 232 is positioned between the high-speed compressor 236 and thehigh-speed turbine 238.

The combustion section 232 may be configured as a deflagrativecombustion section, a rotating detonation combustion section, a pulsedetonation combustion section, and/or other appropriate heat additionsystem. The combustion section 232 may be configured as one or more of arich-burn system or a lean-burn system, or combinations thereof. Instill various examples, the combustion section 232 includes an annularcombustor, a can combustor, a cannular combustor, a trapped vortexcombustor (TVC), or another appropriate combustion system, orcombinations thereof.

The core engine 206 also includes a booster or low-pressure compressorpositioned in flow relationship with the high-pressure compressor 236.The low-pressure compressor 242 is rotatably coupled with thelow-pressure turbine 244 via a low-speed shaft 246 to enable thelow-pressure turbine 244 to drive the low-pressure compressor 242. Thelow-speed shaft 246 is also operably connected to the gear assembly 202to provide power to the fan assembly 204, such as described furtherherein.

It should be appreciated that the terms “low” and “high,” or theirrespective comparative degrees (e.g., “lower” and “higher”, whereapplicable), when used with compressor, turbine, shaft, or spoolcomponents, each refer to relative pressures and/or relative speedswithin an engine unless otherwise specified. For example, a “low spool”or “low-speed shaft” defines a component configured to operate at arotational speed, such as a maximum allowable rotational speed, lowerthan a “high spool” or “high-speed shaft” of the engine. Alternatively,unless otherwise specified, the aforementioned terms may be understoodin their superlative degree. For example, a “low turbine” or “low-speedturbine” may refer to the lowest maximum rotational speed turbine withina turbine section, a “low compressor” or “low speed compressor” mayrefer to the lowest maximum rotational speed turbine within a compressorsection, a “high turbine” or “high-speed turbine” may refer to thehighest maximum rotational speed turbine within the turbine section, anda “high compressor” or “high-speed compressor” may refer to the highestmaximum rotational speed compressor within the compressor section.Similarly, the low-speed spool refers to a lower maximum rotationalspeed than the high-speed spool. It should further be appreciated thatthe terms “low” or “high” in such aforementioned regards mayadditionally, or alternatively, be understood as relative to minimumallowable speeds, or minimum or maximum allowable speeds relative tonormal, desired, steady state, etc. operation of the engine.

The compressors and/or turbines disclosed herein can include variousstage counts. As disclosed herein the stage count includes the number ofrotors or blade stages in a particular component (e.g., a compressor orturbine). For example, in some instances, a low-pressure compressor(which can also be referred to as “a booster”) can comprise 1-8 stages,a high-pressure compressor can comprise 8-15 stages, a high-pressureturbine comprises 1-2 stages, and/or a low-pressure turbine comprises3-7 stages (including exactly 3, 4, or 5 stages). For example, incertain examples, an engine can comprise a one stage low-pressurecompressor, an 11 stage high-pressure compressor, a two stagehigh-pressure turbine, and a 7 stage low-pressure turbine. As anotherexample, an engine can comprise a three stage low-pressure compressor, a10 stage high-pressure compressor, a two stage high-pressure turbine,and a 7 stage low-pressure turbine. As another example, an engine cancomprise a three stage low-pressure compressor, a 10 stage high-pressurecompressor, a two stage high-pressure turbine, and a three stagelow-pressure turbine. As another example, an engine can comprise a fourstage low-pressure compressor, a 10 stage high-pressure compressor, aone stage high-pressure turbine, and a three stage low-pressure turbine.As another example, an engine can comprise a three stage low-pressurecompressor, a 10 stage high-pressure compressor, a two stagehigh-pressure turbine, and a four stage low-pressure turbine. As anotherexample, an engine can comprise a four stage low-pressure compressor, a10 stage high-pressure compressor, a one stage high-pressure turbine,and a four stage low-pressure turbine. In other examples, an engine cancomprise a 1-3 stage low-pressure compressor, an 8-11 stagehigh-pressure compressor, a 1-2 stage high-pressure turbine, and a 3-5stage low-pressure turbine. In some examples, an engine can beconfigured without a low-pressure compressor.

In some examples, a low-pressure turbine is a counter-rotatinglow-pressure turbine comprising inner blade stages and outer bladestages. The inner blade stages extend radially outwardly from an innershaft, and the outer blade stages extend radially inwardly from an outerdrum. In particular examples, the counter-rotating low-pressure turbinecomprises three inner blade stages and three outer blade stages, whichcan collectively be referred to as a six stage low-pressure turbine. Inother examples, the counter-rotating low-pressure turbine comprises fourinner blade stages and three outer blade stages, which can collectivelybe referred to as a seven stage low-pressure turbine.

As discussed in more detail below, the core engine 206 includes the gearassembly 202 that is configured to transfer power from the turbinesection 234 and reduce an output rotational speed at the fan assembly204 relative to the low-speed turbine 244. Examples of the gear assembly202 depicted and described herein can allow for gear ratios suitable forlarge diameter unducted fans (e.g., gear ratios of 4.1-14.0, 4.5-14.0,and/or 6.0-14.0). Additionally, examples of the gear assembly 202provided herein may be suitable within the radial or diametricalconstraints of the core engine 206 within the outer casing 214.

Various gearbox configurations are depicted schematically in FIGS. 9-13. These gearboxes can be used in any of the engines disclosed herein,including the engine 200. Additional details regarding the gearboxes areprovided below.

Engine 200 also includes a vane assembly 210 comprising a plurality ofvanes 212 disposed around engine centerline axis 220. Each vane 212 hasa root 248 and a tip 250, and a span defined therebetween. Vanes 212 canbe arranged in a variety of manners. In some examples, they are not allequidistant from the rotating assembly.

In some examples, vanes 212 are mounted to a stationary frame and do notrotate relative to the engine centerline axis 220 but may include amechanism for adjusting their orientation relative to their axis 254and/or relative to the fan blades 208. For reference purposes, FIG. 2depicts a forward direction denoted with arrow F, which in turn definesthe forward and aft portions of the system.

As depicted in FIG. 2 , the fan assembly 204 is located forward of thecore engine 106 with the exhaust 256 located aft of core engine 206 in a“puller” configuration. Other configurations are possible andcontemplated as within the scope of the present disclosure, such as whatmay be termed a “pusher” configuration where the engine core is locatedforward of the fan assembly. The selection of “puller” or “pusher”configurations may be made in concert with the selection of mountingorientations with respect to the airframe of the intended aircraftapplication, and some may be structurally or operationally advantageousdepending upon whether the mounting location and orientation arewing-mounted, fuselage-mounted, or tail-mounted configurations.

Left- or right-handed engine configurations, useful for certaininstallations in reducing the impact of multi-engine torque upon anaircraft, can be achieved by mirroring the airfoils (e.g., 208, 212)such that the fan assembly 204 rotates clockwise for one propulsionsystem and counterclockwise for the other propulsion system.Alternatively, an optional reversing gearbox can be provided to permit acommon gas turbine core and low-pressure turbine to be used to rotatethe fan blades either clockwise or counterclockwise, i.e., to provideeither left- or right-handed configurations, as desired, such as toprovide a pair of oppositely-rotating engine assemblies can be providedfor certain aircraft installations while eliminating the need to haveinternal engine parts designed for opposite rotation directions.

The engine 200 also includes the gear assembly 202 which includes a gearset for decreasing the rotational speed of the fan assembly 204 relativeto the low-speed turbine 244. In operation, the rotating fan blades 208are driven by the low-speed turbine 244 via gear assembly 202 such thatthe fan blades 208 rotate around the engine centerline axis 220 andgenerate thrust to propel the engine 200, and hence an aircraft on whichit is mounted, in the forward direction F.

In some examples, a gear ratio of the input rotational speed to theoutput rotational speed is greater than 4.1. In particular examples, thegear ratio is within a range of 4.1-14.0, within a range of 4.5-14.0, orwithin a range of 6.0-14.0. In certain examples, the gear ratio iswithin a range of 4.5-12 or within a range of 6.0-11.0. As such, in someexamples, the fan assembly can be configured to rotate at a rotationalspeed of 700-1500 rpm at a cruise flight condition, while the powerturbine (e.g., the low-pressure turbine) is configured to rotate at arotational speed of 5,000-10,000 rpm at a cruise flight condition. Inparticular examples, the fan assembly can be configured to rotate at arotational speed of 850-1350 rpm at a cruise flight condition, while thepower turbine is configured to rotate at a rotational speed of5,500-9,500 rpm a cruise flight condition.

It may be desirable that either or both of the fan blades 208 or thevanes 212 to incorporate a pitch change mechanism such that the bladescan be rotated with respect to an axis of pitch rotation (annotated as228 and 254, respectively) either independently or in conjunction withone another. Such pitch change can be utilized to vary thrust and/orswirl effects under various operating conditions, including to provide athrust reversing feature which may be useful in certain operatingconditions such as upon landing an aircraft.

Vanes 212 can be sized, shaped, and configured to impart a counteractingswirl to the fluid so that in a downstream direction aft of both fanblades 208 and vanes 212 the fluid has a greatly reduced degree ofswirl, which translates to an increased level of induced efficiency.Vanes 212 may have a shorter span than fan blades 208, as shown in FIG.2 . For example, vanes 212 may have a span that is at least 50% of aspan of fan blades 208. In some examples, the span of the vanes can bethe same or longer than the span as fan blades 208, if desired. Vanes212 may be attached to an aircraft structure associated with the engine200, as shown in FIG. 2 , or another aircraft structure such as a wing,pylon, or fuselage. Vanes 212 may be fewer or greater in number than, orthe same in number as, the number of fan blades 208. In some examples,the number of vanes 212 are greater than two, or greater than four, innumber. Fan blades 208 may be sized, shaped, and contoured with thedesired blade loading in mind.

In the example shown in FIG. 2 , an annular 360-degree inlet 258 islocated between the fan assembly 204 and the vane assembly 210 andprovides a path for incoming atmospheric air to enter the core engine206 radially inwardly of at least a portion of the vane assembly 210.Such a location may be advantageous for a variety of reasons, includingmanagement of icing performance as well as protecting the inlet 258 fromvarious objects and materials as may be encountered in operation.

In the example of FIG. 2 , in addition to the open rotor or unducted fanassembly 204 with its plurality of fan blades 208, an optional ductedfan assembly 260 is included behind fan assembly 204, such that theengine 200 includes both a ducted and an unducted fan which both serveto generate thrust through the movement of air at atmospherictemperature without passage through the core engine 206. The ducted fanassembly 260 is shown at about the same axial location as the vane 212,and radially inward of the root 248 of the vane 212. Alternatively, theducted fan assembly 260 may be between the vane 212 and core duct 262 orbe farther forward of the vane 212. The ducted fan assembly 260 may bedriven by the low-pressure turbine 244, or by any other suitable sourceof rotation, and may serve as the first stage of the low-pressurecompressor 242 or may be operated separately. Air entering the inlet 258flows through an inlet duct 264 and then is divided such that a portionflows through a core duct 262 and a portion flows through a fan duct266. Fan duct 266 may incorporate heat exchangers 268 and exhausts tothe atmosphere through an independent fixed or variable nozzle 270 aftof the vane assembly 210 at the aft end of the fan cowl 252 and outsideof the engine core cowl 272. Air flowing through the fan duct 266 thus“bypasses” the core of the engine and does not pass through the core.

Thus, in the example, engine 200 includes an unducted fan formed by thefan blades 208, followed by the ducted fan assembly 260, which directsairflow into two concentric or non-concentric ducts 262 and 266, therebyforming a three-stream engine architecture with three paths for airwhich passes through the fan assembly 204.

In the example shown in FIG. 2 , a slidable, moveable, and/ortranslatable plug nozzle 274 with an actuator may be included in orderto vary the exit area of the nozzle 270. A plug nozzle is typically anannular, symmetrical device that regulates the open area of an exit suchas a fan stream or core stream by axial movement of the nozzle such thatthe gap between the nozzle surface and a stationary structure, such asadjacent walls of a duct, varies in a scheduled fashion thereby reducingor increasing a space for airflow through the duct. Other suitablenozzle designs may be employed as well, including those incorporatingthrust reversing functionality. Such an adjustable, moveable nozzle maybe designed to operate in concert with other systems such as VBV's,VSV's, or blade pitch mechanisms and may be designed with failure modessuch as fully-open, fully-closed, or intermediate positions so that thenozzle 270 has a consistent “home” position to which it returns in theevent of any system failure, which may prevent commands from reachingthe nozzle 270 and/or its actuator.

In some examples, a mixing device 276 can be included in a region aft ofa core nozzle 278 to aid in mixing the fan stream and the core stream toimprove acoustic performance by directing core stream outward and fanstream inward.

Since the engine 200 shown in FIG. 2 includes both an open rotor fanassembly 204 and a ducted fan assembly 260, the thrust output of bothand the work split between them can be tailored to achieve specificthrust, fuel burn, thermal management, and/or acoustic signatureobjectives which may be superior to those of a typical ducted fan gasturbine propulsion assembly of comparable thrust class. The ducted fanassembly 260, by lessening the proportion of the thrust required to beprovided by the unducted fan assembly 104, may permit a reduction in theoverall fan diameter of the unducted fan assembly and thereby providefor installation flexibility and reduced weight.

Operationally, the engine 200 may include a control system that managesthe loading of the respective open and ducted fans, as well aspotentially the exit area of the variable fan nozzle, to providedifferent thrust, noise, cooling capacity, and other performancecharacteristics for various portions of the flight envelope and variousoperational conditions associated with aircraft operation. For example,in climb mode the ducted fan may operate at maximum pressure ratiothere-by maximizing the thrust capability of stream, while in cruisemode, the ducted fan may operate a lower pressure ratio, raising overallefficiency through reliance on thrust from the unducted fan. Nozzleactuation modulates the ducted fan operating line and overall engine fanpressure ratio independent of total engine airflow.

The ducted fan stream flowing through fan duct 266 may include one ormore heat exchangers 268 for removing heat from various fluids used inengine operation (such as an air-cooled oil cooler (ACOC), cooledcooling air (CCA), etc.). The heat exchangers 268 may take advantage ofthe integration into the fan duct 266 with reduced performance penalties(such as fuel efficiency and thrust) compared with traditional ductedfan architectures, due to not impacting the primary source of thrustwhich is, in this case, the unducted fan stream. Heat exchangers maycool fluids such as gearbox oil, engine sump oil, thermal transportfluids such as supercritical fluids or commercially availablesingle-phase or two-phase fluids (supercritical CO2, EGV, Slither 800,liquid metals, etc.), engine bleed air, etc. Heat exchangers may also bemade up of different segments or passages that cool different workingfluids, such as an ACOC paired with a fuel cooler. Heat exchangers 268may be incorporated into a thermal management system which provides forthermal transport via a heat exchange fluid flowing through a network toremove heat from a source and transport it to a heat exchanger.

The fan duct 266 also provides other advantages in terms of reducednacelle drag, enabling a more aggressive nacelle close-out, improvedcore stream particle separation, and inclement weather operation. Byexhausting the fan duct flow over the core cowl, aids in energizing theboundary layer and enabling the option of a steeper nacelle close outangle between the maximum dimension of the engine core cowl 272 and theexhaust 256. The close-out angle is normally limited by air flowseparation, but boundary layer energization by air from the fan duct 266exhausting over the core cowl reduces air flow separation. This yields ashorter, lighter structure with less frictional surface drag.

The fan assembly and/or vane assembly can be shrouded or unshrouded (asshown in FIGS. 1 and 2 ). Although not shown, an optional annular shroudor duct can be coupled to the vane assembly 210 and located distallyfrom the engine centerline axis 220 relative to the vanes 212. Inaddition to the noise reduction benefit, the duct may provide improvedvibratory response and structural integrity of the vanes 212 by couplingthem into an assembly forming an annular ring or one or morecircumferential sectors, i.e., segments forming portions of an annularring linking two or more of the vanes 212. The duct may also allow thepitch of the vanes to be varied more easily. For example, FIGS. 3-4 ,discussed in more detail below, disclose examples in which both the fanassembly and vane assembly are shrouded.

Although depicted above as an unshrouded or open rotor engine in theexamples depicted above, it should be appreciated that aspects of thedisclosure provided herein may be applied to shrouded or ducted engines,partially ducted engines, aft-fan engines, or other turbomachineconfigurations, including those for marine, industrial, oraero-propulsion systems. Certain aspects of the disclosure may beapplicable to turbofan, turboprop, or turboshaft engines. However, itshould be appreciated that certain aspects of the disclosure may addressissues that may be particular to unshrouded or open rotor engines, suchas, but not limited to, issues related to gear ratios, fan diameter, fanspeed, length (L) of the engine, maximum diameter of the core engine(D_(core)) of the engine, L/D_(core) of the engine, desired cruisealtitude, and/or desired operating cruise speed, or combinationsthereof.

FIG. 3 is a schematic cross-sectional view of a gas turbine engine inaccordance with an example of the present disclosure. More particularly,for the example of FIG. 3 , the gas turbine engine is a high-bypassturbofan jet engine 300, referred to herein as “turbofan engine 300.” Asshown in FIG. 3 , the turbofan engine 300 defines an axial direction A(extending parallel to a longitudinal centerline 302 provided forreference) and a radial direction R (extending perpendicular to theaxial direction A). In general, the turbofan 300 includes a fan section304 and a core engine 306 disposed downstream from the fan section 304.The engine 300 also includes a gear assembly or power gear box 336having a plurality of gears for coupling a gas turbine shaft to a fanshaft. The position of the power gear box 336 is not limited to that asshown in the example of turbofan 300. For example, the position of thepower gear box 336 may vary along the axial direction A.

The exemplary core engine 306 depicted generally includes asubstantially tubular outer casing 308 that defines an annular inlet310. The outer casing 308 encases, in serial flow relationship, acompressor section including a booster or low-pressure (LP) compressor312 and a high-pressure (HP) compressor 314; a combustion section 316; aturbine section including a high-pressure (HP) turbine 318 and alow-pressure (LP) turbine 320; and a jet exhaust nozzle section 322. Ahigh-pressure (HP) shaft or spool 324 drivingly connects the HP turbine318 to the HP compressor 314. A low-pressure (LP) shaft or spool 326drivingly connects the LP turbine 320 to the LP compressor 312.Additionally, the compressor section, combustion section 316, andturbine section together define at least in part a core air flowpath 327extending therethrough.

A gear assembly of the present disclosure is compatible with standardfans, variable pitch fans, or other configurations. For the exampledepicted, the fan section 304 may include a variable pitch fan 328having a plurality of fan blades 330 coupled to a disk 332 in aspaced-apart manner. As depicted, the fan blades 330 extend outwardlyfrom disk 332 generally along the radial direction R. Each fan blade 330is rotatable relative to the disk 332 about a pitch axis P by virtue ofthe fan blades 330 being operatively coupled to a suitable actuationmember 334 configured to collectively vary the pitch of the fan blades330. The fan blades 330, disk 332, and actuation member 334 are togetherrotatable about the longitudinal axis 302 by LP shaft 326 across a gearassembly or power gearbox 336. A gear assembly 336 may enable a speedchange between a first shaft, e.g., LP shaft 326, and a second shaft,e.g., LP compressor shaft and/or fan shaft. For example, in someinstances, the gear assembly 336 may be disposed in an arrangementbetween a first shaft and a second shaft such as to reduce an outputspeed from one shaft to another shaft.

More generally, the gear assembly 336 can be placed anywhere along theaxial direction A to decouple the speed of two shafts, whenever it isconvenient to do so from a component efficiency point of view, e.g.,faster LP turbine and slower fan and LP compressor or faster LP turbineand LP compressor and slower fan.

The gear assembly 336 (which can also be referred to as “a gearbox”)can, in some examples, comprise a gear ratio of less than or equal tofour. For example, the gearbox 336 can comprise a gear ratio within arange of 2.0-4.0, 2.0-2.9, 3.2-4.0, 3.25-3.75, etc. In one particularexample, the gearbox 336 can comprise a gear ratio of 3.5.

Referring still to the example of FIG. 3 , the disk 332 is covered byrotatable front nacelle 338 aerodynamically contoured to promote airflowthrough the plurality of fan blades 330. Additionally, the exemplary fansection 304 includes an annular fan casing or outer nacelle 340 thatcircumferentially surrounds the fan 328 and/or at least a portion of thecore engine 306. The nacelle 340 is, for the example depicted, supportedrelative to the core engine 306 by a plurality ofcircumferentially-spaced outlet guide vanes 342. Additionally, adownstream section 344 of the nacelle 340 extends over an outer portionof the core engine 306 so as to define a bypass airflow passage 346therebetween.

During operation of the turbofan engine 300, a volume of air 348 entersthe turbofan 300 through an associated inlet 350 of the nacelle 340and/or fan section 304. As the volume of air 348 passes across the fanblades 330, a first portion of the air 348 as indicated by arrows 352 isdirected or routed into the bypass airflow passage 346 and a secondportion of the air 348 as indicated by arrow 354 is directed or routedinto the LP compressor 312. The ratio between the first portion of air352 and the second portion of air 354 is commonly known as a bypassratio. The pressure of the second portion of air 354 is then increasedas it is routed through the high-pressure (HP) compressor 314 and intothe combustion section 316, where it is mixed with fuel and burned toprovide combustion gases 356.

The combustion gases 356 are routed through the HP turbine 318 where aportion of thermal and/or kinetic energy from the combustion gases 356is extracted via sequential stages of HP turbine stator vanes 358 thatare coupled to the outer casing 308 and HP turbine rotor blades 360(e.g., two stage) that are coupled to the HP shaft or spool 324, thuscausing the HP shaft or spool 324 to rotate, thereby supportingoperation of the HP compressor 314. The combustion gases 356 are thenrouted through the LP turbine 320 where a second portion of thermal andkinetic energy is extracted from the combustion gases 356 via sequentialstages of LP turbine stator vanes 362 that are coupled to the outercasing 308 and LP turbine rotor blades 364 (e.g., four stages) that arecoupled to the LP shaft or spool 326, thus causing the LP shaft or spool326 to rotate, thereby supporting operation of the LP compressor 312and/or rotation of the fan 328.

It should be noted that a high-pressure turbine (e.g., the HP turbine318) can, in some examples, comprise one or two rotating blade stagesand that a low-pressure turbine (e.g., LP turbine 320) can, in someinstances, comprise three, four, five, six, or seven rotating bladestages.

The combustion gases 356 are subsequently routed through the jet exhaustnozzle section 322 of the core engine 306 to provide propulsive thrust.Simultaneously, the pressure of the first portion of air 352 issubstantially increased as the first portion of air 352 is routedthrough the bypass airflow passage 346 before it is exhausted from a fannozzle exhaust section 366 of the turbofan 300, also providingpropulsive thrust. The HP turbine 318, the LP turbine 320, and the jetexhaust nozzle section 322 at least partially define a hot gas path 368for routing the combustion gases 356 through the core engine 306.

For example, FIG. 4 is a cross-sectional schematic illustration of anexample of an engine 400 that includes a gear assembly 402 incombination with a ducted fan assembly 404 and a core engine 406.However, unlike the open rotor configuration of the engine 200, the fanassembly 404 and its fan blades 408 are contained within an annular fancase 480 (which can also be referred to as “a nacelle”) and the vaneassembly 410 and the vanes 412 extend radially between the fan cowl 452(and/or the engine core cowl 472) and the inner surface of the fan case480. As discussed above, the gear assemblies disclosed herein canprovide for increased gear ratios for a fixed gear envelope (e.g., withthe same size ring gear), or alternatively, a smaller diameter ring gearmay be used to achieve the same gear ratios.

The core engine 400 comprises a compressor section 430, a combustorsection 432, and a turbine section 434. The compressor section 430 caninclude a high-pressure compressor 436 and a booster or a low-pressurecompressor 442. The turbine section 434 can include a high-pressureturbine 438 (e.g., one stage) and a low-pressure turbine 444 (e.g.,three stage). The low-pressure compressor 442 is positioned forward ofand in flow relationship with the high-pressure compressor 436. Thelow-pressure compressor 442 is rotatably coupled with the low-pressureturbine 444 via a low-speed shaft 446 to enable the low-pressure turbine444 to drive the low-pressure compressor 442 (and a ducted fan 460). Thelow-speed shaft 446 is also operably connected to the gear assembly 402to provide power to the fan assembly 404. The high-pressure compressor436 is rotatably coupled with the high-pressure turbine 438 via ahigh-speed shaft 440 to enable the high-pressure turbine 438 to drivethe high-pressure compressor 436.

It should be noted that a high-pressure turbine (e.g., the high-pressureturbine 438) can, in some examples, comprise one or two stages and thata low-pressure turbine (e.g., the low-pressure turbine 444) can, in someinstances, comprise three, four, five, or six rotating blade stages.

In some examples, the engine 400 can comprise a pitch change mechanism482 coupled to the fan assembly 404 and configured to vary the pitch ofthe fan blades 408. In certain examples, the pitch change mechanism 482can be a linear actuated pitch change mechanism.

In some examples, the engine 400 can comprise a variable fan nozzle.Operationally, the engine 400 may include a control system that managesthe loading of the fan, as well as potentially the exit area of thevariable fan nozzle, to provide different thrust, noise, coolingcapacity and other performance characteristics for various portions ofthe flight envelope and various operational conditions associated withaircraft operation. For example, nozzle actuation modulates the fanoperating line and overall engine fan pressure ratio independent oftotal engine airflow.

In some examples, an engine (e.g., the engine 100, the engine 200,and/or the engine 400) can comprise a counter-rotating low-pressureturbine. For example, FIGS. 5-6 depict schematic cross-sectionalillustrations of counter-rotating low-pressure turbines. In particular,FIG. 5 depicts a counter-rotating turbine 500, and FIG. 6 depicts acounter-rotating turbine 600. The counter-rotating turbines compriseinner blade stages and outer blade stages arranged in an alternatinginner-outer configuration. In other words, the counter-rotating turbinesdo not comprise stator vanes disposed between the blade stages.

Referring to FIG. 5 , the counter-rotating turbine 500 comprises aplurality of inner blade stages 502 and a plurality of outer bladestages 504. More specifically, the counter-rotating turbine 500 includesthree inner blades stages 502 that are coupled to and extend radiallyoutwardly from an inner shaft 506 (which can also be referred to as “arotor”) and three outer blade stages 504 that are coupled to extendradially inwardly from an outer shaft 508 (which can also be referred toas “a drum”). In this manner, the counter-rotating turbine 500 can beconsidered a six stage turbine because there are six total rotatingblade rows. It should be noted, however, that the counter-rotatingturbine 500 would be considered to have three LPT stages, N, forpurposes of determining the overall engine efficiency rating describedbelow because there are three inner blade stages 502.

Referring to FIG. 6 , the counter-rotating turbine 600 comprises aplurality of inner blade stages 602 and a plurality of outer bladestages 604. More specifically, the counter-rotating turbine 600 includesfour inner blades stages 602 that are coupled to and extend radiallyoutwardly from an inner shaft 606 and three outer blade stages 604 thatare coupled to extend radially inwardly from an outer shaft 608. In thismanner, the counter-rotating turbine 600 can be considered a seven stageturbine because there are seven total rotating blade rows. It should benoted, however, that the counter-rotating turbine 600 would beconsidered to have four LPT stages, N, for purposes of determining theoverall engine efficiency rating described below because there are fourinner blade stages 602.

According to some examples, there is a turbomachine characterized byboth a high gear ratio and a high power gearbox. A high gear ratiogearbox means a gearbox with a gear ratio of above about 4:1. Examplesof a high power gearbox include a gearbox adapted for transmitting powergreater than 7 MW with output spool speed above, e.g., 1000 rpm, agearbox adapted for transmitting power greater than 15 MW with outputspool speed of about 1100 rpm, and a gearbox adapted for transmittingpower greater than transmitting 22 MW with output spool speed of about3500 rpm.

Each of the examples of turbomachines disclosed herein can utilize ahigh gear ratio gearbox. Adopting a gearbox with a high gear ratiopresents unique challenges. One such challenge is determining the amountof oil that would need to circulate through the gearbox duringoperation, i.e., the high gear ratio gearbox's oil flow rate. The oildemand is significant when the engine requires a high gear ratiogearbox. Moreover, the estimated amount of oil flow for the high gearratio gearbox is not well informed by, or capable of being estimatedfrom, oil flow rates for an existing serviced engine. Starting from thisbasis, the oil flow demands were calculated for the different engineconfigurations contemplated and disclosed herein, by consideration ofthe different features and performance characteristics, e.g., pitch linevelocity and constants differentiating one gearbox configuration fromanother. The high gear ratio gearbox architectures considered includethose described and disclosed herein (e.g., FIGS. 9-13 and theaccompanying text, infra). These efforts accordingly involved factoringin specific characteristics of the gearboxes and the power transmissionrequirements for the gearbox to estimate the oil flow rates.

During the process of developing the aforementioned examples ofturbomachines incorporating a high gear ratio gearbox, it was determinedthat a good approximation of the high gear ratio gearbox oil flow ratemay be made using only a relatively few engine parameters. Thisdevelopment is based on, among other things, the recognition that an oilflow rate through a gearbox is related to the expected power loss whentransmitting power across a gearbox. From this initial recognition andother developments that were the by-product of studying severaldifferent engine configurations that included a power gearbox (includingthe configurations disclosed herein), it was determined that a goodapproximation to the high gear ratio gearbox oil flow rate could be madebased on a relationship among the turbomachine's gearbox gear ratio, netthrust, and fan diameter. This relationship is referred to herein as “agearbox efficiency rating.”

The gearbox efficiency rating is quite beneficial. For example, with thegearbox efficiency rating having provided the engine oil flowrequirements one can also estimate, for purposes of system integration,the type of oil-related secondary systems (e.g., sump, oil circuit, heatsinks, etc.) that would be included to support proper functioning of theselected high gear ratio gearbox; and/or to provide guidance on whethera particular engine architecture is beneficial or not, without requiringan entire team to complete the tedious and time-consuming process ofdeveloping a new gearbox from scratch. Therefore, the gearbox efficiencyrating can improve the process of developing a turbomachine engine,which can ultimately result in an improved turbomachine. Values for thegearbox efficiency rating identify key engine requirements affecting theoverall architecture. An engine architecture based, at least in part, onthis value, can enable early optimization of major engine components,thereby benefiting the overall architecture. By basing an engine designon a gearbox efficiency rating, it is more likely to find the optimizedarchitecture than versus a design on experiment. The GER enablesdiscovery of a better design for this reason, rather than relying onchance that the optimal solution is found from a design of experimentsinvolving a large number of variables whose interrelationships are notclearly known or understood.

As indicated, the gearbox efficiency rating is a relationship based on aturbomachine's fan diameter (D), net thrust (T), and gear ratio of ahigh gear ratio gearbox. The gear efficiency rating, valid for gearratios between about 4:1 and 14:1, may be expressed as

${Q\left( \frac{D^{1.56}}{T} \right)}^{1.53},$

where Q is measured at an inlet of the gearbox in gallons per minute ata max takeoff condition, D is measured in inches, and T is measured inpounds force at the max takeoff condition. In this manner, the gearboxefficiency rating defines a specific turbomachine engine configuration.

As used herein “net thrust” (T) equals the change of momentum of thebypass airflow plus the change of momentum of the core airflow and theburned fuel. Or stated another way, T=W_(byp)(V_(byp)−V₀)+(W_(core)+W_(fuel)) V_(core)−W_(core) V₀, where W_(byp) isthe mass flow rate of air of the bypass airflow, V_(byp) is the velocityof the bypass airflow, V₀ is the flight velocity, W_(core) is the massflow rate of air of the core airflow, W_(fuel) is burned the mass flowrate of the burned fuel, and V_(core) is the velocity of the coreairflow.

As indicated earlier, turbomachine engines, such as the turbofan engines100, 200, 300, 400, comprise many variables and factors that affecttheir performance and/or operation. The interplay between the variouscomponents can make it particularly difficult to develop or select onecomponent, especially when each of the components is at a differentstage of completion. For example, one or more components may be nearlycomplete, yet one or more other components may be in an initial orpreliminary phase where only one (or a few) parameters is known. Also,each component is subject to change often more than once over thedevelopment period, which can often last for many years (e.g., 5-15years). These complex and intricate individual and collectivedevelopment processes can be cumbersome and inefficient. For at leastthese reasons, there is a need for devices and methods that can providea good estimate of, not only the basic configuration or sizing needed toachieve the desired performance benefits, but also to reflect thepenalties or accommodations in other areas in order to realize thedesired benefits. This leads to an improved, more optimally designedengine.

According to another aspect of the disclosure, the gearbox efficiencyrating may additionally provide a particularly useful indication of theefficiency and effectiveness of the engine during initial development,e.g., as a tool to accept or reject a particular configuration. Thus,the gearbox efficiency rating can be used, for example, to guide gearboxdevelopment. For example, the gearbox efficiency rating can be used toquickly and accurately determine the size of the gearbox that issuitable for a particular engine without requiring an individual or teamto complete the tedious and time-consuming process of developing thegearbox from scratch. Therefore, the gearbox efficiency rating can alsoimprove the process of developing a turbomachine engine.

As further explained below, the inventors also discovered thatmodification of the gearbox efficiency rating accounting for the numberof rotating low-pressure turbine stages, referred to as overall engineefficiency rating, could also improve the overall engine architecture.One way in which the overall engine efficiency rating improves enginearchitecture is that balances several engine parameters to provide awell-balanced and efficient engine. The overall engine efficiency ratingcan also, for example, aid in the process of developing a turbomachineengine. The overall gearbox efficiency rating can be particularly usefulfor geared turbofan engines comprising a gear ratio that is less than orequal to 4.0 and/or for ducted, geared turbofan engines.

FIGS. 7A-8 illustrate exemplary ranges and/or values for gear efficiencyrating. FIGS. 7A-7C disclose exemplary ranges of gear efficiency ratingwith respect to various gear ratios. FIG. 8 discloses the gear ratio,oil flow, fan diameter, net thrust, and gearbox efficiency ratings formultiple exemplary turbomachine engines.

In some examples, the gearbox efficiency rating of a turbomachine engineis within a range of about 0.10-1.8 or 0.19-1.8 or 0.10-1.01. In certainexamples, the gearbox efficiency rating is within a range of about0.25-0.55 or about 0.29-0.51. FIG. 8 provides the gear efficiency ratingof several exemplary engines.

Since a gearbox is used as a speed reducer or increaser in transmittingpower from component to component, gearbox efficiency is of primaryimportance. Various dynamic issues invariably will arise during theextended operation of the power gearbox. Accordingly, the ability of thebearings to tolerate and mitigate these dynamic issues can improve thecapacity, life, and reliability of the power gearbox and thereby lowerthe frequency of the engine maintenance. Additionally, providing properlubrication and cooling to the bearings and/or other gearbox componentsis necessary to maximize the life and load capacity gearbox. Thus, anychanges to the engine architecture (e.g., fan diameter) must notadversely affect proper lubrication and cooling to the gearbox. Thegearbox efficiency rating takes this into account and provides an engineconfiguration with proper oil flow rate. As such, the gearbox efficiencyrating can, for example, provide an engine with improved gearboxefficiency and/or increased longevity.

In some examples, the oil flow rate Q is within a range of about 5-55gallons per minute. In certain examples, the oil flow rate Q is within arange of about 5.5-25 gallons per minute. In other examples, the oilflow rate Q is within a range of about 25-55 gallons per minute. Inother examples, the oil flow rate Q is within a range of about 25-40gallons per minute. In other examples, the oil flow rate Q is within arange of about 20-30 gallons per minute. In other examples, the oil flowrate Q is within a range of about 25-35 gallons per minute. FIGS. 8 and20 also provides the oil flow rates of several exemplary engines.

As noted above, the oil flow rate Q is measured at an inlet of thegearbox in gallons per minute at a max takeoff condition. The inlet ofthe gearbox is the location at which the oil enters the gearbox from theoil supply line. As used herein “a max takeoff condition” meanssea-level elevation, standard pressure, extreme hot day temperature, anda flight velocity of up to about 0.25 Mach.

As used herein, the term “extreme hot day temperature” means the extremehot day temperature specified for a particular engine. This can includethe extreme hot day temperature used for engine certification. Extremehot day temperature can additionally or alternatively includetemperatures of about 130-140° F.

In some examples, the fan diameter D is about 120-216 inches. In certainexamples, the fan diameter D is about 120-192 inches. FIG. 8 alsoprovides the fan diameter of several exemplary engines.

In some examples, the net thrust T of the engine is within a range ofabout 10,000-100,000 pounds force. In particular examples, the netthrust T of the engine is within a range of about 12,000-30,000 poundsforce. FIG. 8 also provides the net thrust of several exemplary engines.

In some examples, the gearbox efficiency rating of a turbomachine enginecan be configured in relation to the gear ratio (GR) of the gearbox. Forexample, in certain instances, a turbomachine engine can be configuredsuch that the gearbox efficiency rating is greater than 0.015 (GR^(1.4))and less than 0.034 (GR^(1.5)), as depicted in FIG. 7A. In otherexamples, a turbomachine engine can be configured such that the gearboxefficiency rating is greater than 0.02625 (GR^(1.4)) and less than 0.042(GR^(1.4)).

For example, FIG. 8 depicts several exemplary engines with gearboxefficiency ratings that satisfy these relationships. Engine 1 is aturbomachine engine comprising a gearbox with a gear ratio of 10.5:1 anda gearbox efficiency rating within a range of 0.40-1.16, specifically1.02. Engine 2, Engine 3, and Engine 4 are turbomachine enginescomprising gearboxes with a gear ratio of 7:1 and the gearbox efficiencyratings within a range of 0.23-0.63, that is 0.51, 0.42, and 0.41,respectively. Engine 5 is a turbomachine engine comprising a gearboxwith a gear ratio of 5.1:1 and a gearbox efficiency rating within arange of 0.15-0.39, specifically 0.29. Engine 6 is a turbomachine enginecomprising a gearbox with a gear ratio of 4.1:1 and a gearbox efficiencyrating within a range of 0.11-0.28, specifically 0.21. Engines 7-19provide additional examples with specific gearbox efficiency ratings.Ranges for the gearbox efficiency ratings of Engines 7-19 can bedetermined using the equations above and/or the charts of FIGS. 7A-7C.

As another example, a turbomachine engine comprising a gearbox with agear ratio of 4.5:1 can be configured such that the gearbox efficiencyrating is within a range of 0.12-0.32. As another example, aturbomachine engine comprising a gearbox with a gear ratio of 6:1 can beconfigured such that the gearbox efficiency rating is within a range of0.18-0.50. As another example, a turbomachine engine comprising agearbox with a gear ratio of 9:1 can be configured such that the gearboxefficiency rating is within a range of 0.33-0.92. As another example, aturbomachine engine comprising a gearbox with a gear ratio of 11:1 canbe configured such that the gearbox efficiency rating within a range of0.43-1.24. As another example, a turbomachine engine comprising agearbox with a gear ratio of 12:1 can be configured such that thegearbox efficiency rating within a range of 0.49-1.41. As yet anotherexample, a turbomachine engine comprising a gearbox with a gear ratio of14:1 can be configured such that the gearbox efficiency rating is withina range of 0.60-1.78.

In some instances, a turbomachine engine can comprise a gearbox with agear ratio of 5-6, 7-8, 9-10, 11-12, or 13-14. In other instances, aturbomachine engine can comprise a gearbox with a gear ratio of 5-7,8-10, 11-13. In yet other examples, a turbomachine engine can comprise agearbox with a gear ratio of 7-10 or 11-14. Below is a table withseveral exemplary gearbox efficiency ratings with respect to severalexemplary gear ratios.

Gear Ratio Gearbox Efficiency Rating 4.1-6.9 0.10-0.62 7.0-9.9 0.22-1.0610.0-12.9 0.37-1.56 13.0-14.0 0.54-1.8 

In some examples, a turbomachine engine can be configured such that thegearbox efficiency rating is greater than 0.023 (GR^(1.5)) and less than0.034 (GR^(1.5)), as depicted in FIG. 7B. In particular instances, thegearbox efficiency rating can be about 0.0275 (GR^(1.5)). Theseconfigurations can be particularly advantageous, for example, withengines comprising an epicyclic gearbox (e.g., star and/or planetconfiguration).

Gear Ratio Gearbox Efficiency Rating 4.1-6.9 0.19-0.62 7.0-9.9 0.43-1.0610.0-12.9 0.73-1.58 13.0-14.0 1.08-1.8 

In other examples, a turbomachine engine can be configured such that thegearbox efficiency rating is greater than 0.015 (GR^(1.4)) and less than0.025 (GR^(1.4)), as depicted in FIG. 7C. In particular instances, thegearbox efficiency rating can be about 0.02 (GR^(1.4)). Theseconfigurations can be particularly advantageous, for example, withengines comprising a non-epicyclic gearbox (e.g., compound gearboxes).

Gear Ratio Gearbox Efficiency Rating 4.1-6.9 0.10-0.37 7.0-9.9 0.23-0.6210.0-12.9 0.38-0.90 13.0-14.0 0.54-1.01

It should be noted gearbox efficiency rating values disclosed herein areapproximate values. Accordingly, the disclosed gearbox efficiency ratingvalues include values within five percent of the listed values.

As noted above, the gearbox efficiency rating can define a specificengine configuration and/or can be used when developing a gearbox for aturbomachine engine. For example, in some instances, the gearboxefficiency rating can be used to determine the size and/or oil flow rateof a gearbox. Assuming that a desired gear ratio of the gearbox isknown, along with the fan diameter, and the net thrust of the engine,the gearbox efficiency ratings depicted in the charts of FIG. 7A-7C canbe used to determine an acceptable oil flow rate. In some examples, theequation below can be used to determine an acceptable range of oil flowrates (Q) for the gearbox. The determined oil flow rate Q can be used,for example, to aid in the configuration of the gearbox, thereby leadingto an improved gearbox and the overall engine. In some instances, one ormore other parameters (e.g., the gearbox efficiency rating) can also aidin the configuration of the gearbox.

$\frac{0.015\left( {GR}^{1.4} \right)}{\left( \frac{D^{1.56}}{T} \right)^{1.53}} < Q < \frac{0.034\left( {GR}^{1.5} \right)}{\left( \frac{D^{1.56}}{T} \right)^{1.53}}$

For example, a gearbox for a turbomachine engine can be configured usingthe following exemplary method. With reference to FIG. 8 , Engine 1comprises an unducted fan and can be configured similar to the engine200. Engine 1 comprises a fan diameter of 188.6 inches and a net thrustof 25,503 pounds force at a max takeoff condition. Engine 1 furthercomprises a five stage low-pressure turbine. The desired gear ratio forthe gearbox of Engine 1 is about 10.5:1. Based on this information, theoil flow rate Q of the gearbox of Engine 1 should be about 8-24 gallonsper minute at a max takeoff condition.

FIG. 9 schematically depicts a gearbox 700 that can be used, forexample, with Engine 1. The gearbox 700 comprises a two-stage starconfiguration.

The first stage of the gearbox 700 includes a first-stage sun gear 702,a first-stage carrier 704 housing a plurality of first-stage star gears,and a first-stage ring gear 706. The first-stage sun gear 702 can becoupled to a low-speed shaft 708, which in turn is coupled to thelow-pressure turbine of Engine 1. The first-stage sun gear 702 can meshwith the first-stage star gears, which mesh with the first-stage ringgear. The first-stage carrier 704 can be fixed from rotation by asupport member 710.

The second stage of the gearbox 700 includes a second-stage sun gear712, a second-stage carrier 714 housing a plurality of second-stage stargears, and a second-stage ring gear 716. The second-stage sun gear 712can be coupled to a shaft 718 which in turn is coupled to thefirst-stage ring gear 706. The second-stage carrier 714 can be fixedfrom rotation by a support member 720. The second-stage ring gear 716can be coupled to a fan shaft 722.

In some examples, each stage of the gearbox 700 can comprise five stargears. In other examples, the gearbox 700 can comprise fewer or morethan five star gears in each stage. In some examples, the first-stagecarrier can comprise a different number of star gears than thesecond-stage carrier. For example, the first carrier can comprise fivestar gears, and the second-stage carrier can comprise three star gears,or vice versa.

Based on the configuration of the gearbox 700 and the calculated oilflow rate of 8-24 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 700 can comprise a radius R₁. The size ofthe gearbox, including the radius R₁, can be configured such that theoil flow rate at the inlet of the gearbox 700 at a max takeoff conditionis about 8-24 gallons per minute or about 16-24 gallons per minute(e.g., 20.9 gpm). In some examples, the radius R₁ of the gearbox 700 canbe about 16-19 inches. In other examples, the radius R₁ of the gearbox700 can be about 22-24 inches. In other examples, the radius R₁ of thegearbox 700 can be smaller than 16 inches or larger than 24 inches.

As another example, Engine 2 (FIG. 8 ) comprises an unducted fan and canbe configured similar to the engine 200. Engine 3 comprises a fandiameter of 188.6 inches and a net thrust of 25,000 pounds force at amax takeoff condition. Engine 2 further comprises a 3-7 stagelow-pressure turbine. The desired gear ratio for the gearbox of Engine 2is about 7:1. Based on this information, oil flow rate Q of the gearboxof Engine 2 should be about 4-13 gallons per minute or about 8-13gallons per minute (e.g., 10.06) at a max takeoff condition.

FIG. 10 schematically depicts a gearbox 800 that can be used, forexample, with Engine 2. The gearbox 800 comprises a single-stage starconfiguration. The gearbox 800 includes a sun gear 802, a carrier 804housing a plurality of star gears (e.g., 3-5 star gears), and a ringgear 806. The sun gear 802 can mesh with the star gears, and the stargears can mesh with the ring gear 806. The sun gear 802 can be coupledto a low-speed shaft 808, which in turn is coupled to the low-pressureturbine of Engine 2. The carrier 804 can be fixed from rotation by asupport member 810. The ring gear 806 can be coupled to a fan shaft 812.

Based on the configuration of the gearbox 800 and the calculated oilflow rate of 4-13 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 800 can comprise a radius R₂. The size ofthe gearbox, including the radius R₂, can be configured such that theoil flow rate at the inlet of the gearbox 800 at a max takeoff conditionis 7-13 gallons per minute (e.g., 10.1 gpm). In some examples, theradius R₂ of the gearbox 800 can be about 18-23 inches. In otherexamples, the radius R₂ of the gearbox 700 can be smaller than 18 inchesor larger than 23 inches.

As another example, Engine 3 (FIG. 8 ) comprises an unducted fan and canbe configured similar to the engine 200. Engine 3 comprises a fandiameter of 142.8 inches and a net thrust of 12,500 pounds force at amax takeoff condition. Engine 3 further comprises a 3-7 stagelow-pressure turbine. The desired gear ratio for the gearbox of Engine 3is about 7:1. Based on this information, oil flow rate Q of the gearboxof Engine 3 should be about 3-9 gallons per minute or about 5-9 gallonsper minute (e.g., 6 gpm) at a max takeoff condition.

FIG. 11 schematically depicts a gearbox 900 that can be used, forexample, with Engine 3. The gearbox 800 comprises a single-stage starconfiguration. The gearbox 900 includes a sun gear 902, a carrier 904housing a plurality of star gears (e.g., 3-5 star gears), and a ringgear 906. The sun gear 902 can mesh with the star gears, and the stargears can mesh with the ring gear 906. The sun gear 902 can be coupledto a low-speed shaft 908, which in turn is coupled to the low-pressureturbine of Engine 3. The carrier 904 can be fixed from rotation by asupport member 910. The ring gear 906 can be coupled to a fan shaft 912.

Based on the configuration of the gearbox 900 and the calculated oilflow rate of 5-9 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 900 can comprise a radius R₃. The size ofthe gearbox, including the radius R₃, can be configured such that theoil flow rate at the inlet of the gearbox 900 at a max takeoff conditionis 3-9 gallons per minute (e.g., 6 gpm). In some examples, the radius R₃of the gearbox 900 can be about 10-13 inches. In other examples, theradius R₃ of the gearbox 900 can be smaller than 10 inches or largerthan 13 inches.

Engine 4 comprises an unducted fan and can be configured similar to theengine 200. Engine 4 comprises a fan diameter of 188.4 inches and a netthrust of 25,000 pounds force at a max takeoff condition. Engine 4further comprises a counter-rotating low-pressure turbine (e.g., similarto the counter-rotating turbine 500 or the counter-rotating turbine600). The desired gear ratio for the gearbox of Engine 4 is about 7:1.Based on this information, oil flow rate Q of the gearbox of Engine 4should be about 4-13 gallons per minute or about 7-13 gallons per minute(e.g., 8.1 gpm) at a max takeoff condition.

FIG. 12 schematically depicts a gearbox 1000 that can be used, forexample, with Engine 4. The gearbox 1000 comprises a two-stageconfiguration in which the first stage is a star configuration and thesecond stage is a planet configuration.

The first stage of the gearbox 1000 includes a first-stage sun gear1002, a first-stage star carrier 1004 comprising a plurality offirst-stage star gears (e.g., 3-5 star gears), and a first-stage ringgear 1006. The first-stage sun gear 1002 can mesh with the first-stagestar gears, and the first-stage star gears can mesh with the first-stagering gear 1006. The first-stage sun gear 1002 can be coupled to ahigher-speed shaft 1008 of the low spool, which in turn is coupled tothe inner blades of the low-pressure turbine of Engine 4. Thefirst-stage star carrier 1004 can be fixed from rotation by a supportmember 1010.

The second stage of the gearbox 1000 includes a second-stage sun gear1012, a second-stage planet carrier 1014 comprising a plurality ofsecond-stage planet gears (e.g., 3-5 planet gears), and a second-stagering gear 1016. The second-stage sun gear 1012 can mesh with thesecond-stage planet gears. The second-stage planet carrier 1014 can becoupled to the first-stage ring gear 1006. The second-stage sun gear1012 can be coupled to a lower-speed shaft 1018 of the low spool, whichin turn is coupled to the outer blades of the low-pressure turbine ofEngine 4. The second-stage planet carrier 1014 can be coupled to thefirst-stage ring gear 1006. The second-stage planet carrier 1014 canalso be coupled to a fan shaft 1020. The second-stage ring gear 1016 canbe fixed from rotation by a support member 1022.

In some examples, each stage of the gearbox 1000 can comprise threestar/planet gears. In other examples, the gearbox 1000 can comprisefewer or more than three star/planet gears in each stage. In someexamples, the first-stage carrier can comprise a different number ofstar gears than the second-stage carrier has planet gears. For example,the first-carrier can comprise five star gears, and the second-stagecarrier can comprise three planet gears, or vice versa.

Since the first stage of the gearbox 1000 is coupled to the higher-speedshaft 1008 of the low spool and the second stage of the gearbox 1000 iscoupled to the lower-speed shaft 1018 of the low spool, the gear ratioof the first stage of the gearbox 1000 can be greater than the gearratio of the second stage of the gearbox. For example, in certainconfigurations, the first stage of the gearbox can comprise a gear ratioof 4.1-14, and the second stage of the gearbox can comprise a gear ratiothat is less than the gear ratio of the first stage of the gearbox. Inparticular examples, the first stage of the gearbox can comprise a gearratio of 7, and the second stage of the gearbox can comprise a gearratio of 6.

In some examples, an engine comprising the gearbox 1000 can beconfigured such that the higher-speed shaft 1008 provides about 50% ofthe power to the gearbox 1000 and the lower-speed shaft 1018 providesabout 50% of the power to the gearbox 1000. In other examples, an enginecomprising the gearbox 1000 can be configured such that the higher-speedshaft 1008 provides about 60% of the power to the gearbox 1000 and thelower-speed shaft 1018 provides about 40% of the power to the gearbox1000.

Based on the configuration of the gearbox 1000 and the calculated oilflow rate of 4-13 gallons per minute, which is based on the gearboxefficiency rating, the gearbox 1000 can comprise a radius R₄. The sizeof the gearbox, including the radius R₄, can be configured such that theoil flow rate at the inlet of the gearbox 1000 at a max takeoffcondition is 7-13 gallons per minute (e.g., 8.1 gpm). In some examples,the radius R₄ of the gearbox 1000 can be about 18-22 inches. In otherexamples, the radius R₄ of the gearbox 700 can be smaller than 18 inchesor larger than 22 inches.

Thus, as illustrated by the examples disclosed herein, a gearboxefficiency rating can characterize or define a specific engine and/orgearbox configuration. As such, turbomachine engines can be quickly andaccurately configured by utilizing the gearbox efficiency rating and/orits related parameters. In this manner, the gearbox efficiency ratingdisclosed herein provides one or more significant advantages over knownturbomachine engines and/or known methods of developing turbomachineengines.

FIG. 13 depicts a gearbox 1100 that can be used, for example, with theengines disclosed herein (e.g., the engines 100, 200, 400). The gearbox1100 is configured as a compound star gearbox. The gearbox 1100comprises a sun gear 1102 and a star carrier 1104, which includes aplurality of compound star gears having one or more first portions 1106and one or more second portions 1108. The gearbox 1100 further comprisesa ring gear 1110. The sun gear 1102 can also mesh with the firstportions 1106 of the star gears. The star carrier can be fixed fromrotation via a support member 1114. The second portions 1108 of the stargears can mesh with the ring gear 1110. The sun gear 1102 can be coupledto a low-pressure turbine via the turbine shaft 1112. The ring gear 1110can be coupled to a fan shaft 1116.

The gear assemblies shown and described herein can be used with anysuitable engine. For example, although FIG. 4 shows an optional ductedfan and optional fan duct (similar to that shown in FIG. 2 ), it shouldbe understood that such gear assemblies can be used with other ductedturbofan engines (e.g., the engine 300) and/or other open rotor enginesthat do not have one or more of such structures.

Configurations of the gear assemblies depicted and described herein mayprovide for gear ratios and arrangements that fit within the L/D_(core)constraints of the disclosed engines. In certain examples, the gearassemblies depicted and described in regard to FIGS. 9-13 allow for gearratios and arrangements providing for rotational speed of the fanassembly corresponding to one or more ranges of cruise altitude and/orcruise speed provided above.

Various configurations of the gear assembly provided herein may allowfor gear ratios of up to 14:1. Still various examples of the gearassemblies provided herein may allow for gear ratios of at least 4.1:1or 4.5:1. Still yet various examples of the gear assemblies providedherein allow for gear ratios of 6:1 to 12:1. Other examples can have agear ratio within a range of 2.0-4.0. FIGS. 8 and 20 also provide thegear ratio of several exemplary engines. It should be appreciated thatexamples of the gear assemblies provided herein may allow for large gearratios and within constraints such as, but not limited to, length of theengine, maximum diameter (D_(core)) of the engine 100, cruise altitudeof up to 65,000 ft, and/or operating cruise speed of up to Mach 0.85, orcombinations thereof. The disclosed gear assemblies may alternatively beconfigured to provide a gear ratio that is within a range of 2.0-4.0.

Various exemplary gear assemblies are shown and described herein. Thesegear assemblies may be utilized with any of the exemplary engines and/orany other suitable engine for which such gear assemblies may bedesirable. In such a manner, it will be appreciated that the gearassemblies disclosed herein may generally be operable with an enginehaving a rotating element with a plurality of rotor blades and aturbomachine having a turbine and a shaft rotatable with the turbine.With such an engine, the rotating element (e.g., fan assembly 104) maybe driven by the shaft (e.g., low-speed shaft 146) of the turbomachinethrough the gear assembly.

Although the exemplary gear assemblies shown are mounted at a forwardlocation (e.g., forward from the combustor and/or the low-pressurecompressor), in other examples, the gear assemblies described herein canbe mounted at an aft location (e.g., aft of the combustor and/or thelow-pressure turbine).

Portions of a lubricant system 1200 are depicted schematically in FIG.14 . The lubrication system 1200 can be a component of the turbomachineengines disclosed herein and/or can be coupled to the various gearboxesdisclosed herein. For example, FIG. 1 schematically illustrates thelubricant system coupled to the turbomachine engine 100 and the gearbox102. A series of lubricant conduits 1203 can interconnect multipleelements of the lubricant system 1200 and/or engine components, therebyproviding for provision or circulation of the lubricant throughout thelubricant system and any engine components coupled thereto (e.g., agearbox, bearing compartments, etc.).

It should be understood that the organization of the lubricant system1200 as shown is by way of example only to illustrate an exemplarysystem for a turbomachine engine for circulating lubricant for purposessuch as lubrication or heat transfer. Any organization for the lubricantsystem 1200 is contemplated, with or without the elements as shown,and/or including additional elements interconnected by any necessaryconduit system.

Referring again to FIG. 14 , the lubricant system 1200 includes alubricant reservoir 1202 configured to store a coolant or lubricant,including organic or mineral oils, synthetic oils, or fuel, or mixturesor combinations thereof. A supply line 1204 and a scavenge line 1206 arefluidly coupled to the reservoir 1202 and collectively form a lubricantcircuit to which the reservoir 1202 and component 1210 (e.g., a gearbox)can be fluidly coupled. The component 1210 can be supplied withlubrication by way of a fluid coupling with the supply line 1204 and canreturn the supplied lubricant to the reservoir 1202 by fluidly couplingto the scavenge line 1206. More specifically, a component supply line1211 can be fluidly coupled between the supply line 1204 and thecomponent 1210. It is further contemplated that multiple types oflubricant can be provided in other lines not explicitly shown but arenonetheless included in the lubricant system 1200.

Optionally, at least one heat exchanger 1205 can be included in thelubricant system 1200. The heat exchanger 1205 can include afuel/lubricant (fuel-to-lubricant) heat exchanger, an oil/lubricant heatexchanger, an air-cooled oil cooler, and/or other means for exchangingheat. For example, a fuel/lubricant heat exchanger can be used to heator cool engine fuel with lubricant passing through the heat exchanger.In another example, a lubricant/oil heat exchanger can be used to heator cool additional lubricants passing within the turbomachine engine,fluidly separate from the lubricant passing along the lubricant system1200. Such a lubricant/oil heat exchanger can also include aservo/lubricant heat exchanger. Optionally, a second heat exchanger (notshown) can be provided along the exterior of the core engine, downstreamof the outlet guide vane assembly. The second heat exchanger can be anair/lubricant heat exchanger, for example, adapted to convectively coollubricant in the lubricant system 1200 utilizing the airflow passingthrough an outlet guide vane assembly of the turbomachine engine.

A pump 1208 can be provided in the lubricant system 1200 to aid inrecirculating lubricant from the reservoir 1202 to the component 1210via the supply line 1204. For example, the pump 1208 can be driven by arotating component of the turbine engine 10, such as a high-pressureshaft or a low-pressure shaft of a turbomachine engine.

Lubricant can be recovered from the component 1210 by way of thescavenge line 1206 and returned to the reservoir 1202. In theillustrated example, the pump 1208 is illustrated along the supply line1204 downstream of the reservoir 1202. The pump 1208 can be located inany suitable position within the lubricant system 1200, including alongthe scavenge line 1206 upstream of the reservoir 1202. In addition,while not shown, multiple pumps can be provided in the lubricant system1200.

In some examples, a bypass line 1212 can be fluidly coupled to thesupply line 1204 and scavenge line 1206 in a manner that bypasses thecomponent 1210. In such examples, a bypass valve 1215 is fluidly coupledto the supply line 1204, component supply line 1211, and bypass line1212. The bypass valve 1215 is configured to control a flow of lubricantthrough at least one of the component supply line 1211 or the bypassline 1212. The bypass valve 1215 can include any suitable valveincluding, but not limited to, a differential thermal valve, rotaryvalve, flow control valve, and/or pressure safety valve. In someexamples, a plurality of bypass valves can be provided.

During operation, a supply flow 1220 can move from the reservoir 1202,through the supply line 1204, and to the bypass valve 1215. A componentinput flow 1222 can move from the bypass valve 1215 through thecomponent supply line 1211 to an inlet of the component 1210. A scavengeflow 1224 can move lubricant from an outlet of the component 1210through the scavenge line 1206 and back to the reservoir 1202.Optionally, a bypass flow 1226 can move from the bypass valve 1215through the bypass line 1212 and to the scavenge line 1206. The bypassflow 1226 can mix with the scavenge flow 1224 and define a return flow1228 moving toward the lubricant reservoir 1202.

In one example where no bypass flow exists, it is contemplated that thesupply flow 1220 can be the same as the component input flow 1222 andthat the scavenge flow 1224 can be the same as the return flow 1228. Inanother example where the bypass flow 1226 has a nonzero flow rate, thesupply flow 1220 can be divided at the bypass valve 1215 into thecomponent input flow 1222 and bypass flow 1226. It will also beunderstood that additional components, valves, sensors, or conduit linescan be provided in the lubricant system 1200, and that the example shownin FIG. 14 is simplified with a single component 1210 for purposes ofillustration.

The lubricant system 1200 can further include at least one sensingposition at which at least one lubricant parameter can be sensed ordetected. The at least one lubricant parameter can include, but is notlimited to, a flow rate, a temperature, a pressure, a viscosity, achemical composition of the lubricant, or the like. In the illustratedexample, a first sensing position 1216 is located in the supply line1204 upstream of the component 1210, and a second sensing position 1218is located in the scavenge line 1206 downstream of the component 1210.

In one example, the bypass valve 1215 can be in the form of adifferential thermal valve configured to sense or detect at least onelubricant parameter in the form of a temperature of the lubricant. Insuch a case, the fluid coupling of the bypass valve 1215 to the firstand second sensing positions 1216, 1218 can provide for bypass valve1215 sensing or detecting the lubricant temperature at the sensingpositions 1216, 18 as lubricant flows to or from the bypass valve 1215.The bypass valve 1215 can be configured to control the component inputflow 1222 or the bypass flow 1226 based on the sensed or detectedtemperature.

It is contemplated that the bypass valve 1215, supply line 1204, andbypass line 1212 can at least partially define a closed-loop controlsystem for the component 1210. As used herein, a “closed-loop controlsystem” will refer to a system having mechanical or electroniccomponents that can automatically regulate, adjust, modify, or control asystem variable without manual input or other human interaction. Suchclosed-loop control systems can include sensing components to sense ordetect parameters related to the desired variable to be controlled, andthe sensed or detected parameters can be utilized as feedback in a“closed loop” manner to change the system variable and alter the sensedor detected parameters back toward a target state. In the example of thelubricant system 1200, the bypass valve 1215 (e.g., mechanical orelectrical component) can sense a parameter, such as a lubricantparameter (e.g., temperature), and automatically adjust a systemvariable, e.g., flow rate to either or both of the bypass line 1212 orcomponent 1210, without need of additional or manual input. In oneexample, the bypass valve can be automatically adjustable orself-adjustable such as a thermal differential bypass valve. In anotherexample, the bypass valve can be operated or actuated via a separatecontroller. It will be understood that a closed-loop control system asdescribed herein can incorporate such a self-adjustable bypass valve ora controllable bypass valve.

Turning to FIG. 15 , a portion of the lubricant system 1200 isillustrated supplying lubricant to a particular component 1210 in theform of a gearbox 1250 within a turbomachine engine. The gearbox can beany of the gearboxes disclosed herein. The gearbox 1250 can include aninput shaft 1252, an output shaft 1254, and a gear assembly 1255. In oneexample, the gear assembly 1255 can be in the form of an epicyclic gearassembly as known in the art having a ring gear, sun gear, and at leastone planet/star gear. An outer housing 1256 can at least partiallysurround the gear assembly 1255 and form a structural support for thegears and bearings therein. Either or both of the input and outputshafts 1252, 1254 can be coupled to the turbomachine engine. In oneexample, the input and output shafts 1252, 1254 can be utilized todecouple the speed of the low-pressure turbine from the low-pressurecompressor and/or the fan, which can, for example, improve engineefficiency.

The supply line 1204 can be fluidly coupled to the gearbox 1250, such asto the gear assembly 1255, to supply lubricant to gears or bearings tothe gearbox 1250 during operation. The scavenge line 1206 can be fluidlycoupled to the gearbox 1250, such as to the gear assembly 1255 or outerhousing 1256, to collect lubricant. The bypass line 1212 can be fluidlycoupled to the bypass valve 1215, supply line 1204, and scavenge line1206 as shown. A return line 1214 can also be fluidly coupled to thebypass valve 1215, such as for directing the return flow 1228 to thelubricant reservoir 1202 for recirculation. While not shown in FIG. 15for brevity, the lubricant reservoir 1202, the heat exchanger 1205,and/or the pump 1208 (FIG. 14 ) can also be fluidly coupled to thegearbox 1250. In this manner, the supply line 1204, bypass line 1212,scavenge line 1206, and return line 1214 can at least partially define arecirculation line 1230 for the lubricant system 1200.

The supply flow 1220 divides at the bypass line into the component inputflow 1222 and the bypass flow 1226. In the example shown, the bypassvalve 1215 is in the form of a differential thermal valve that isfluidly coupled to the first and second sensing positions 1216, 1218.

Lubricant flowing proximate the first and second sensing positions 1216,1218 provides the respective first and second outputs 1241, 1242indicative of the temperature of the lubricant at those sensingpositions 1216, 1218. It will be understood that the supply line 1204 isthermally coupled to the bypass line 1212 and bypass valve 1215 suchthat the temperature of the fluid in the supply line 1204 proximate thefirst sensing position 1216 is approximately the same as fluid in thebypass line 1212 adjacent the bypass valve 1215. Two values being“approximately the same” as used herein will refer to the two values notdiffering by more than a predetermined amount, such as by more than 20%,or by more than 5 degrees, in some examples. In this manner, the bypassvalve 1215 can sense the lubricant temperature in the supply line 1204and scavenge line 1206 via the first and second outputs 1241, 1242. Itcan be appreciated that the bypass line 1212 can form a sensing line forthe valve 1215 to sense the lubricant parameter, such as temperature, atthe first sensing position 1216.

During operation of the turbomachine engine, the lubricant temperaturecan increase within the gearbox 1250, such as due to heat generation ofthe gearbox 1250, and throughout the lubricant system 1200. In oneexample, if a lubricant temperature exceeds a predetermined thresholdtemperature at either sensing position 1216, 1218, the bypass valve 1215can automatically increase the component input flow 1222, e.g., from thesupply line 1204 to the gearbox 1250, by decreasing the bypass flow1226. Such a predetermined threshold temperature can be any suitableoperating temperature for the gearbox 1250, such as about 300° F. insome examples. Increasing the component input flow 1222 can provide forcooling of the gearbox 1250, thereby reducing the lubricant temperaturesensed in the various lines 1204, 1206, 1212, 1214 as lubricantrecirculates through the lubricant system 1200.

In another example, if a temperature difference between the sensingpositions 1216, 1218 exceeds a predetermined threshold temperaturedifference, the bypass valve can automatically increase the componentinput flow 1222 by decreasing the bypass flow 1226. Such a predeterminedthreshold temperature difference can be any suitable operatingtemperature for the gearbox 1250, such as about 70° F., or differing bymore than 30%, in some examples. In yet another example, if atemperature difference between the sensing positions 1216, 1218 is belowthe predetermined threshold temperature difference, the bypass valve canautomatically decrease the component input flow 1222 or increase thebypass flow 1226. In this manner the lubricant system 1200 can providefor the gearbox to operate with a constant temperature differencebetween the supply and scavenge lines 1204, 1206.

Starting from the basis of the gearbox efficiency rating, the inventorsset out to determine whether the gearbox efficiency rating (and/or itscomponents) could be used to aid in the process of developing and/orapply to a geared turbofan engine comprising a relatively low gear ratio(e.g., a gear ratio less than or equal to 4.0—e.g., 2.0-4.0). Afternumerous attempts and analyzing a multitude of engine parameters andengine configurations, the inventors discovered that the gearboxefficiency rating, when taken together with the stage count of thelow-pressure turbine, can in some cases provide an improved engineconfiguration compared to an engine configuration based only on gearboxefficiency rating, particularly for engines comprising a gear ratio lessthan or equal to 4.0 (e.g., 2.0-4.0). More precisely, the inventorsdiscovered an overall engine efficiency rating, which is a relationshipbetween the gearbox (i.e., the oil flow “Q”), the fan (i.e., the fandiameter “D”), the power output (i.e., the net thrust “T”), and thelow-pressure turbine (i.e., the number of LPT stages “N”). The overallengine efficiency rating can in some cases identify a more holisticengine configuration, which can, for example, improve the efficiency ofthe engine. In addition to an improved overall engine configuration, theoverall engine efficiency rating can in some instances be used to guidean engine development process.

The overall engine efficiency, valid for gear ratios within a range of2.0-4.0, is defined as Q(D^(1.56)/T)^(1.53) N², where Q is a gearbox oilflow rate at an inlet of the gearbox measured in gallons per minute at amax takeoff condition, D is a diameter of the fan blades measured ininches, T is a net thrust of the turbomachine engine measured in poundsforce at the max takeoff condition, and N is a number of rotating bladestages of the low-pressure turbine. This newly developed engineparameter can, for example, aid in the process of developing aturbomachine because it considers parameters of a turbofan engine andprovides a good approximation of an engine's overall efficiency early onin development. Values for the overall engine efficiency rating identifykey engine requirements affecting the overall architecture, in a similarmanner as the gearbox efficiency rating discussed earlier. The overallengine efficiency rating however may be a more insightful value toidentify an optimal solution because, in addition to the oil flow, theoverall engine efficiency rating factors in the effects on architecturewhen the number of LPT stages are increased or decreased. When there isan increase in the number of LPT stages the turbine efficiency improves,but there is a weight penalty. It may be necessary to balance the numberof LPT stages against the size of the gearbox, oil flow needs to thegearbox, and/or size of the fan. An engine architecture based, at leastin part, on a value dependent on both the gearbox and LPT, can similarlyenable early optimization of major engine components, thereby benefitingthe overall architecture. By basing an engine design on an overallengine efficiency rating, it is more likely to find the optimizedarchitecture than versus a design of experiment. The overall engineefficiency rating enables improved engine configurations for thisreason, rather than relying on chance that the optimal solution is foundfrom a design of experiments involving a large number of variables whoseinterrelationships are not clearly known or understood.

As noted above, turbomachine engines, such as the turbofan engines 100,200, 300, 400, comprise many variables and factors that affect theirperformance and/or operation. The interplay between the variouscomponents can make it particularly difficult to develop or select onecomponent, especially when each of the components is at a differentstage of completion. For example, one or more components may be nearlycomplete, yet one or more other components may be in an initial orpreliminary phase where only one (or a few) parameters is known. Also,each component is subject to change often more than once over thedevelopment period, which can often last for many years (e.g., 5-15years). These complex, intricate individual and collective developmentprocesses can be cumbersome and inefficient. For at least these reasons,the overall engine efficiency rating can provide a good estimate of, notonly the basic configuration or sizing needed to achieve the desiredperformance benefits, but also to reflect the penalties oraccommodations in other areas in order to realize the desired benefits.

According to another aspect of the disclosure, the overall engineefficiency rating may additionally provide a particularly usefulindication of the efficiency and effectiveness of the engine duringinitial development, e.g., as a tool to accept or reject a particularconfiguration. Thus, the overall engine efficiency rating can be used,for example, to turbofan engine development. For example, the overallengine efficiency rating can be used to quickly and accurately determineparameters (e.g., the size of the gearbox, the number of LPT stages,and/or size of the fan) that are suitable for a particular enginewithout requiring an individual or team to complete the tedious andtime-consuming process of developing the entire engine or a componentfrom scratch. In this manner, the overall engine efficiency rating canalso improve the process of developing a turbomachine engine.

The overall engine efficiency rating can be particularly advantageous indeveloping ducted geared turbofan engines. For example, the overallengine efficiency rating can be utilized for the ducted geared turbofanengines 300 and 400 disclosed herein.

It should be noted that the number of LPT stages (N) of a low-pressureturbine for purposes of the determining the overall engine efficiencyrating of a turbofan engine defined as the number of rotating bladestages (or rotors) of the low-pressure turbine for a low-pressureturbine that includes blade (rotor) and vane (stator) rows. When thelow-pressure turbine is a counter-rotating turbine (i.e., without vanesbetween adjacent rotating blade rows), the number of LPT stages (N) isthe number of inner blade stages (as opposed to outer blade stages ortotal rotating stages). For example, referring to FIG. 5 , thecounter-rotating low-pressure turbine 500 has three inner blade stages502, and thus “N” equals three when determining the overall engineefficiency for an engine comprising the counter-rotating low-pressureturbine 500. For example, referring to FIG. 6 , the counter-rotatinglow-pressure turbine 600 has four inner blade stages 602, and thus “N”equals four when determining the overall engine efficiency for an enginecomprising the counter-rotating low-pressure turbine 600.

In some examples, the overall engine efficiency rating can be greaterthan or equal to 0.1GR^(1.5) and less than or equal to GR^(1.5), whereGR is the gear ratio. For example, FIGS. 16A-18B depict various rangesof the overall engine efficiency rating and the gear ratio that satisfythis relationship.

FIG. 16A depicts overall engine efficiency rating within a range of0.57-8.0 for gear ratios within a range of 3.2-4.0, where the overallengine efficiency rating is greater than or equal to 0.1GR^(1.5) andless than or equal to GR^(1.5). This range for overall engine efficiencyrating and/or gear ratios may be particularly advantageous whenconfiguring an engine to meet today and future demands, including fuelefficiency and power.

FIG. 16B depicts a subrange of the overall engine efficiency rating ofFIG. 16A. Specifically, FIG. 16B depicts an overall engine efficiencyrating of 0.57-3.0 for gear ratios of 3.2-4.0, where the overall engineefficiency rating greater than or equal to 0.1GR^(1.5) and less than orequal to 3.0. Configuring an engine within the subrange depicted in FIG.16B can, for example, provide a relatively light and/or efficientengine. As another example, an engine comprising an overall engineefficiency rating within the subrange of FIG. 16B can be relativelycompact, which can be advantageous when sizing/space is at a premium.

FIG. 16C depicts another subrange of the overall engine efficiencyrating of FIG. 16A. Particularly, FIG. 16C depicts an overall engineefficiency rating within a range of 3.0-8.0 for gear ratios of 3.2-4.0,where the overall engine efficiency rating is greater than or equal to3.0 and less than or equal to GR^(1.5). Configuring an engine within thesubrange depicted in FIG. 16C can, for example, provides a less costlyand/or more durable engine. In particular examples, engines within thesubrange depicted in FIG. 16C can have relatively higher oil flows ratesthan the engines within the subrange depicted in FIG. 16B. This can,among other things, reduce gearbox temperatures. As a result, lessexpensive materials can be used within the gearbox. Additionally (oralternatively), the durability of the gearbox can be improved and/orservice intervals can be extended.

FIG. 16D depicts a subrange of the overall engine efficiency rating ofFIG. 16A. More precisely, FIG. 16D depicts overall engine efficiencyrating within a range of 0.59-7.3 for gear ratios within a range of3.25-3.75, where the overall engine efficiency rating is greater than orequal to 0.1GR^(1.5) and less than or equal to GR^(1.5).

FIG. 17A depicts overall engine efficiency rating within a range of0.28-4.9 for gear ratios within a range of 2.0-2.9, where the overallengine efficiency rating is greater than or equal to 0.1GR^(1.5) andless than or equal to GR^(1.5). The range of overall engine efficiencyrating depicted in FIG. 17A can, for example, be advantageous forengines comprising a counter-rotating low-pressure turbine orconfigurations where lower turbine speeds would produce a more efficientsystem due to aerodynamic or mechanical constraints.

FIG. 17B depicts a subrange of the overall engine efficiency rating ofFIG. 17A. FIG. 17B depicts overall engine efficiency rating within arange of 0.28-3.9 for gear ratios within a range of 2.0-2.5, where theoverall engine efficiency rating is be greater than or equal to0.1GR^(1.5) and less than or equal to GR^(1.5).

FIG. 17C depicts a subrange of the overall engine efficiency rating ofFIG. 17A. In particular, FIG. 17C depicts overall engine efficiencyrating within a range of 0.9-2.1 for gear ratios within a range of2.0-2.5, where the overall engine efficiency rating is greater than orequal to 0.1GR^(1.5) and less than or equal to GR^(1.5).

The overall engine efficiency ratings depicted in FIGS. 17A-17C can, insome examples, be particularly advantageous for engines comprising acounter-rotating low-pressure turbine. Specifically, the range depictedin FIG. 17C can be particularly well suited for engines comprising acounter-rotating low-pressure turbine or configurations where lowerturbine speeds would produce a more efficient system due to aerodynamicor mechanical constraints.

FIG. 18A depicts overall engine efficiency rating within a range of1.9-8.0 for gear ratios within a range of 2.0-4.0, where the overallengine efficiency rating is greater than or equal to 1.9 and less thanor equal to GR^(1.5). The range depicted in FIG. 18A can, in someinstances, produce an engine that is less costly and/or more durablethan other geared engines having a gear ratio between 2.0 and 4.0.

FIG. 18B depicts a subrange of the overall engine efficiency rating ofFIG. 18A. Specifically, FIG. 18B depicts overall engine efficiencyrating within a range of 1.9-3.1 for gear ratios within a range of2.0-4.0, where the overall engine efficiency rating is greater than orequal to 1.9 and less than or equal to 3.1. The subrange range depictedin FIG. 18B can, for example, produce an engine that is well balancedand efficient. An engine configured with an overall engine efficiencyrating within the subrange of FIG. 18B can, in some examples, provide agearbox with oil flow rates that keep gearbox temperatures low enough tomake the gearbox durable, while also not having excessive weight.

FIG. 19A depicts overall engine efficiency rating within a range of0.98-5.6 for gear ratios within a range of 2.0-4.0, where the overallengine efficiency rating is greater than or equal to 0.35GR^(1.5) andless than or equal to 0.7GR^(1.5). The subrange range depicted in FIG.19A can, for example, produce an engine that is well balanced,efficient, and cost effective.

FIG. 19B depicts a subrange of the overall engine efficiency rating ofFIG. 19A. Particularly, FIG. 19B depicts overall engine efficiencyrating within a range of 0.98-2.77 for gear ratios within a range of2.0-2.5, where the overall engine efficiency rating is greater than orequal to 0.35GR^(1.5) and less than or equal to 0.7GR^(1.5). The rangedepicted in FIG. 19B can be particularly well suited for enginescomprising a counter-rotating low-pressure turbine or configurationswhere lower turbine speeds would produce a more efficient system due toaerodynamic or mechanical constraints.

FIG. 19C depicts a subrange of the overall engine efficiency rating ofFIG. 19A. Specifically, FIG. 19C depicts overall engine efficiencyrating within a range of 2.0-5.6 for gear ratios within a range of3.2-4.0, where the overall engine efficiency rating is greater than orequal to 0.35GR^(1.5) and less than or equal to 0.7GR^(1.5).

FIG. 20 is a table disclosing several exemplary engines and variousengine parameters that fall within one or more of the overall engineefficiency rating ranges disclosed in FIGS. 16A-19C. The enginesdisclosed in FIG. 20 can, for example, provide a both a fuel efficientand powerful engine.

The engines disclosed herein and comprising the overall engineefficiency rating and/or the gear ratio ranges can, in some instances,comprise a three, a four, or a five stage low-pressure turbine.

This written description uses examples to disclose the technology,including the best mode, and also to enable any person skilled in theart to practice the disclosed technology, including making and using anydevices or systems and performing any incorporated methods. Thepatentable scope of the disclosed technology is defined by the claims,and may include other examples that occur to those skilled in the art.Such other examples are intended to be within the scope of the claims ifthey include structural elements that do not differ from the literallanguage of the claims, or if they include equivalent structuralelements with insubstantial differences from the literal languages ofthe claims.

Further aspects of the disclosure are provided by the subject matter ofthe following examples:

1. A turbomachine engine comprising: a fan assembly including aplurality of fan blades; a vane assembly including a plurality of vanesdisposed aft of the plurality of fan blades; a core engine including alow-pressure turbine; a gearbox including an input and an output,wherein the input of the gearbox is coupled to the low-pressure turbineof the core engine and comprises a first rotational speed, wherein theoutput of the gearbox is coupled to the fan assembly and has a secondrotational speed, and wherein a gear ratio of the first rotational speedto the second rotational speed is within a range of 3.2-4.0; and anoverall engine efficiency rating of 0.57-8.0, wherein the overall engineefficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

2. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-1.0.

3. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-1.5.

4. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-2.0.

5. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-2.5.

6. The turbomachine engine of any example herein, and particularlyexample 1, wherein the overall engine efficiency rating is 0.57-3.0.

7. The turbomachine engine of any example herein, and particularlyexample 1, wherein the overall engine efficiency rating is within arange 0.8-3.0.

8. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-3.5.

9. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-4.0.

10. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-4.5.

11. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-5.0.

12. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-5.5.

13. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-6.0.

14. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-6.5.

15. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-7.0.

16. The turbomachine of any example herein, and particularly example 1,wherein the overall engine efficiency rating is 0.57-7.5.

17. The turbomachine engine of any example herein, and particularlyexample 1, wherein the overall engine efficiency rating is 3.0-8.0.

18. The turbomachine engine of any example herein, and particularly anyone of examples 1-17, wherein the gear ratio is within a range of3.5-4.0.

19. The turbomachine engine of any example herein, and particularly anyone of examples 1-17, wherein the gear ratio is within a range of3.4-3.6.

20. The turbomachine engine of any example herein, and particularly anyone of examples 1-19, wherein Q is within a range of 5-55 gallons perminute.

21. The turbomachine engine of any example herein, and particularly anyone of examples 1-19, wherein Q is within a range of 26-45 gallons perminute.

22. The turbomachine engine of any example herein, and particularly anyone of examples 1-21, wherein D is 80-160 inches.

23. The turbomachine engine of any example herein, and particularly anyone of examples 1-21, wherein D is 90-120 inches.

24. The turbomachine engine of any example herein, and particularly anyone of examples 1-23, wherein T is within a range of 10,000-100,000pounds force.

25. The turbomachine engine of any example herein, and particularly anyone of examples 1-23, wherein T is within a range of 25,000-40,000pounds force.

26. The turbomachine engine of any example herein, and particularly anyone of examples 1-25, wherein the gearbox is an epicyclic gearboxcomprising a sun gear, a plurality of planet gears, and a ring gear,wherein the sun gear is the input, and wherein the ring gear is theoutput.

27. The turbomachine engine of any example herein, and particularly anyone of examples 1-25, wherein the gearbox is an epicyclic gearboxcomprising a sun gear, a plurality of planet gears, and a ring gear,wherein the sun gear is the input, wherein the planet gears are coupledto a planet carrier, and wherein the planet carrier is the output.

28. The turbomachine engine of any example herein, and particularly anyone of examples 1-27, wherein the low-pressure turbine includes exactlythree stages.

29. The turbomachine engine of any example herein, and particularly anyone of examples 1-27, wherein the low-pressure turbine includes exactlyfour stages.

30. The turbomachine engine of any example herein, and particularly anyone of examples 1-27, wherein the low-pressure turbine includes exactlyfive stages.

31. A turbomachine engine comprising: a fan casing; a fan assemblydisposed within the fan case and including a plurality of fan blades; avane assembly disposed within the fan case and including a plurality ofvanes disposed aft of the plurality of fan blades; a core engineincluding a low-pressure compressor, a high-pressure compressor, acombustor, a high-pressure turbine, and a low-pressure turbine, whereinthe high-pressure turbine is coupled to the high-pressure compressor viaa high-speed shaft, and wherein the low-pressure turbine is coupled tothe low-speed compressor via a low-speed shaft; a gearbox including aninput and an output, wherein the input of the gearbox is coupled to thelow-speed shaft and comprises a first rotational speed, wherein theoutput of the gearbox is coupled to the fan assembly and has a secondrotational speed, and wherein a gear ratio of the first rotational speedto the second rotational speed is within a range of 3.25-3.75; and anoverall engine efficiency rating of 0.59-7.3, wherein the overall engineefficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotor stages of the low-pressureturbine.

32. The turbomachine engine of any example herein, and particularlyexample 31, wherein Q is within a range of 15-35 gallons per minute,wherein D is 80-150 inches, and wherein T is within a range of25,000-40,000 pounds force.

33. The turbomachine engine of any example herein, and particularlyeither example 31 or example 32, wherein the gearbox is an epicyclicgearbox comprising a star gear configuration.

34. The turbomachine engine of any example herein, and particularlyeither example 31 or example 32, wherein the gearbox is an epicyclicgearbox comprising a planet gear configuration.

35. The turbomachine engine of any example herein, and particularly anyone of examples 31-34, wherein the fan assembly comprises 12-18 fanblades, wherein the low-pressure compressor comprises 1-8 stages,wherein the high-pressure compressor comprises 8-15 stages, wherein thehigh-pressure turbine comprises 1-2 stages, and the low-pressure turbinecomprises 3-5 stages.

36. A turbomachine engine comprising: a ducted fan assembly including aplurality of fan blades; a ducted vane assembly including a plurality ofvanes, wherein the plurality of vanes is configured to receive a firstportion of airflow from the plurality of fan blades; a core engineconfigured to receive a second portion of the airflow from the pluralityof fan blades, wherein the core engine includes a low-pressurecompressor, a high-pressure compressor, a combustor, a high-pressureturbine, and a low-pressure turbine; a gearbox comprising a gear ratiowithin a range of 3.2-4.0; and an overall engine efficiency rating of0.8-3.0.

37. The turbomachine engine of any example herein, and particularlyexample 36, wherein a gearbox oil flow rate at an inlet of the gearboxis within a range of 20-55 gallons per minute at a max takeoffcondition.

38. The turbomachine engine of any example herein, and particularlyeither example 36 or example 37, wherein a diameter of the fan blades is80-144 inches.

39. The turbomachine engine of any example herein, and particularly anyone of examples 36-38, wherein a net thrust of the turbomachine engineis within a range of 25,000-80,000 pounds force at a max takeoffcondition.

40. A turbomachine engine comprising: a nacelle; a fan assembly disposedwithin the nacelle and including a plurality of fan blades arranged in asingle blade row; a vane assembly disposed within the nacelle andincluding a plurality of vanes arranged in a single vane row anddisposed aft of the plurality of fan blades; a core engine including afirst compressor section, a second compressor section, a first turbinesection, and a second turbine section; a first shaft coupling the firstturbine section to the first compressor section; a second shaft couplingthe second turbine section to the second compressor section; a gearboxincluding an input and an output, wherein the input of the gearbox iscoupled to the first shaft and comprises a first rotational speed,wherein the output of the gearbox is coupled to the fan assembly and hasa second rotational speed, which is less than the first rotationalspeed, and wherein a gear ratio of the gearbox is within a range of3.2-4.0; and an overall engine efficiency rating of 0.57-8.0, whereinthe gearbox efficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thefirst turbine section and equals 4.

41. The turbomachine engine of any example herein, and particularlyexample 40, wherein the gearbox is an epicyclic gearbox comprising agear ratio within a range of 3.4-4.0.

42. The turbomachine engine of any example herein, and particularlyeither example 40 or example 41, wherein the overall engine rating iswithin a range of 1.0-3.0.

43. The turbomachine engine of any example herein, and particularlyeither example 40 or example 41, wherein the overall engine rating iswithin a range of 2.0-3.0.

44. The turbomachine engine of any example herein, and particularlyeither example 40 or example 41, wherein the overall engine rating iswithin a range of 2.5-3.0.

45. A turbomachine engine comprising: a fan assembly including aplurality of fan blades; a vane assembly including a plurality of vanesdisposed aft of the plurality of fan blades; a core engine including alow-pressure turbine; a gearbox including an input and an output,wherein the input of the gearbox is coupled to the low-pressure turbineof the core engine and comprises a first rotational speed, wherein theoutput of the gearbox is coupled to the fan assembly and has a secondrotational speed, and wherein a gear ratio (GR) of the first rotationalspeed to the second rotational speed is within a range of 3.2-4.0; andan overall engine efficiency rating greater than 0.1GR^(1.5) and lessthan GR^(1.5), wherein the overall engine efficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

46. The turbomachine engine of any example herein, and particularlyexample 45, wherein the overall engine efficiency rating is greater than0.35GR^(1.5) and less than 0.7GR1.5.

47. The turbomachine engine of any example herein, and particularlyeither example 45 or example 46, wherein the low-pressure turbinecomprises exactly 3 stages.

48. The turbomachine engine of any example herein, and particularlyeither example 45 or example 46, wherein the low-pressure turbinecomprises exactly 4 stages.

49. The turbomachine engine of any example herein, and particularlyeither example 45 or example 46, wherein the low-pressure turbinecomprises exactly 5 stages.

50. A turbomachine engine comprising: a fan assembly including aplurality of fan blades; a vane assembly including a plurality of vanesdisposed aft of the plurality of fan blades; a core engine including alow-pressure turbine; a gearbox including an input and an output,wherein the input of the gearbox is coupled to the low-pressure turbineof the core engine and comprises a first rotational speed, wherein theoutput of the gearbox is coupled to the fan assembly and has a secondrotational speed, and wherein a gear ratio (GR) of the first rotationalspeed to the second rotational speed is within a range of 2.0-2.9; andan overall engine efficiency rating greater than 0.1GR^(1.5) and lessthan GR^(1.5), wherein the overall engine efficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

51. The turbomachine engine of any example herein, and particularlyexample 50, wherein the overall engine efficiency rating is greater than0.35GR^(1.5) and less than 0.7GR^(1.5).

52. The turbomachine engine of any example herein, and particularlyeither example 50 or example 51, wherein the low-pressure turbinecomprises exactly 3 stages.

53. The turbomachine engine of any example herein, and particularlyeither example 50 or example 51, wherein the low-pressure turbinecomprises exactly 4 stages.

54. The turbomachine engine of any example herein, and particularlyeither example 50 or example 51, wherein the low-pressure turbinecomprises exactly 5 stages.

55. The turbomachine engine of any example herein, and particularly anyone of examples 50-54, wherein the gear ratio is within a range of2.0-2.5.

56. A turbomachine engine comprising: a fan assembly including aplurality of fan blades; a vane assembly including a plurality of vanesdisposed aft of the plurality of fan blades; a core engine including alow-pressure turbine; a gearbox including an input and an output,wherein the input of the gearbox is coupled to the low-pressure turbineof the core engine and comprises a first rotational speed, wherein theoutput of the gearbox is coupled to the fan assembly and has a secondrotational speed, and wherein a gear ratio (GR) of the first rotationalspeed to the second rotational speed is within a range of 2.0-4.0; andan overall engine efficiency rating greater than 0.35GR^(1.5) and lessthan 0.7GR^(1.5), wherein the overall engine efficiency rating equals

${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$

wherein Q is a gearbox oil flow rate at an inlet of the gearbox measuredin gallons per minute at a max takeoff condition, wherein D is adiameter of the fan blades measured in inches, wherein T is a net thrustof the turbomachine engine measured in pounds force at the max takeoffcondition, and wherein N is a number of rotating blade stages of thelow-pressure turbine.

57. The turbomachine engine of any example herein, and particularlyexample 56, wherein the gear ratio is within a range of 2.0-2.5.

58. The turbomachine engine of any example herein, and particularlyexample 56, wherein the gear ratio is within a range of 3.2-4.0.

59. The turbomachine engine of any example herein, and particularlyexample 56, wherein the gear ratio is within a range of 3.25-3.75.

60. The turbomachine engine of any example herein, and particularly anyone of examples 56-59, wherein the overall engine efficiency rating isless than 3.0.

61. The turbomachine engine of any example herein, and particularly anyone of examples 56-60, wherein Q is within a range of 5-52 gallons perminute.

62. The turbomachine engine of any example herein, and particularly anyone of examples 56-61, wherein D is 36-144 inches.

63. The turbomachine engine of any example herein, and particularly anyone of examples 56-62, wherein T is within a range of 5,000-80,000pounds force.

64. The turbomachine engine of any example herein, and particularly anyone of examples 56-63, wherein the low-pressure turbine comprises 3, 4,or 5 rotating blade stages.

65. The turbomachine engine of any example herein, and particularly anyone of examples 56-63, wherein the low-pressure turbine comprises 3 or 4rotating blade stages.

66. The turbomachine engine of any example herein, and particularly anyone of examples 56-63, wherein the low-pressure turbine comprisesexactly 4 rotating blade stages.

67. The turbomachine engine of any example herein, wherein the gearboxis an epicyclic gearbox comprising a sun gear, a plurality of planetgears, and a ring gear, wherein the sun gear is the input, and whereinthe ring gear is the output.

68. The turbomachine engine of any example herein, wherein the gearboxis an epicyclic gearbox comprising a sun gear, a plurality of planetgears, and a ring gear, wherein the sun gear is the input, wherein theplanet gears are coupled to a planet carrier, and wherein the planetcarrier is the output.

69. The turbomachine engine of any example herein, wherein the gearboxis a multi-stage gearbox.

70. The turbomachine engine of any example herein, wherein the gearboxis a two-stage gearbox.

71. The turbomachine engine of any example herein, wherein the gearboxis a compound gearbox.

72. The turbomachine engine of any example herein, wherein the fanassembly comprises 8-20 fan blades.

73. The turbomachine engine of any example herein, wherein thelow-pressure compressor comprises 1-4 stages.

74. The turbomachine engine of any example herein, wherein thehigh-pressure compressor comprises 8-11 stages.

75. The turbomachine engine of any example herein, wherein thehigh-pressure turbine comprises 1-2 stages.

76. The turbomachine engine of any example herein, wherein thelow-pressure turbine comprises 3-5 stages.

77. The turbomachine engine of any example herein, wherein thelow-pressure turbine is a counter-rotating low-pressure turbinecomprising inner blade stages and outer blade stages, wherein the innerblade stages extend radially outwardly from an inner shaft, and whereinthe outer blade stages extend radially inwardly from an outer drum.

78. The turbomachine engine of any example herein, wherein thecounter-rotating low-pressure turbine comprises four inner blade stagesand three outer blade stages.

79. The turbomachine engine of any example herein, wherein thecounter-rotating low-pressure turbine comprises three inner blade stagesand three outer blade stages.

80. The turbomachine engine of any example herein, wherein the fanassembly comprises an unducted fan and a ducted fan.

81. The turbomachine engine of any example herein, wherein thelow-pressure compressor comprises 1-2 stages.

82. The turbomachine engine of any example herein, wherein thehigh-pressure compressor comprises 8-10 stages.

83. The turbomachine engine of any example herein, wherein thehigh-pressure turbine comprises two stages.

84. The turbomachine engine of any example herein, wherein thelow-pressure turbine comprises 3-4 stages.

85. The turbofan engine of any example herein, wherein the unducted fanassembly is configured to direct a first portion of airflow to theunducted vane assembly and a second portion of airflow into an inletduct and to the ducted fan assembly, and wherein the ducted fan assemblyis configured to direct the second portion of airflow to a fan duct andto a core duct.

86. The turbomachine engine of any example herein, wherein the gearboxis located forward from the combustor.

87. The turbomachine engine of any example herein, wherein the gearboxis located aft of the combustor.

88. The turbomachine engine of any example herein, wherein the gearboxcomprises one or more compound gears, wherein each compound gearincludes a first portion having a first diameter and a second portionhaving a second diameter, the second diameter being less than the firstdiameter.

89. The turbomachine engine of any example herein, further comprisingone or more pitch change mechanisms coupled to the fan assembly or thevane assembly

1. A turbomachine engine comprising: a fan casing; a fan assemblydisposed within the fan casing and including a plurality of fan blades;a vane assembly disposed within the fan casing and including a pluralityof vanes disposed aft of the plurality of fan blades; a core engineincluding a low-pressure compressor, a high-pressure compressor, acombustor, a high-pressure turbine, and a low-pressure turbine, whereinthe high-pressure turbine is coupled to the high-pressure compressor viaa high-speed shaft, and wherein the low-pressure turbine is coupled tothe low-pressure compressor via a low-speed shaft; a gearbox includingan input and an output, wherein the input of the gearbox is coupled tothe low-speed shaft and comprises a first rotational speed, wherein theoutput of the gearbox is coupled to the fan assembly and has a secondrotational speed, and wherein a gear ratio (GR) of the first rotationalspeed to the second rotational speed is within a range of 2.0-4.0; andan overall engine efficiency rating greater than or equal to0.35GR^(1.5) and less than or equal to 0.7GR^(1.5), wherein the overallengine efficiency rating equals${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$ wherein Q is agearbox oil flow rate at an inlet of the gearbox measured in gallons perminute at a max takeoff condition, wherein D is a diameter of the fanblades measured in inches, wherein T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition, andwherein N is a number of rotor stages of the low-pressure turbine. 2.The turbomachine engine of claim 1, wherein Q is within a range of 15-35gallons per minute, wherein D is 80-150 inches, and wherein T is withina range of 25,000-40,000 pounds force.
 3. The turbomachine engine ofclaim 1, wherein the gearbox is an epicyclic gearbox comprising a stargear configuration.
 4. The turbomachine engine of claim 1, wherein thegearbox is an epicyclic gearbox comprising a planet gear configuration.5. The turbomachine engine of claim 1, wherein the fan assemblycomprises 12-18 fan blades, wherein the low-pressure compressorcomprises 1-5 stages, wherein the high-pressure compressor comprises8-10 stages, wherein the high-pressure turbine comprises 2 stages, andthe low-pressure turbine comprises 3-4 stages.
 6. A turbomachine enginecomprising: a ducted fan assembly including a plurality of fan blades; aducted vane assembly including a plurality of vanes, wherein theplurality of vanes is configured to receive a first portion of airflowfrom the plurality of fan blades; a core engine configured to receive asecond portion of the airflow from the plurality of fan blades, whereinthe core engine includes a low-pressure compressor, a high-pressurecompressor, a combustor, a high-pressure turbine, and a low-pressureturbine; a gearbox comprising a gear ratio (GR) within a range of2.0-2.5; and an overall engine efficiency rating greater than or equalto 0.35GR^(1.5) and less than or equal to 0.7GR^(1.5), wherein theoverall engine efficiency rating equals${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$ wherein Q is agearbox oil flow rate at an inlet of the gearbox measured in gallons perminute at a max takeoff condition, wherein D is a diameter of the fanblades measured in inches, wherein T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition, andwherein N is a number of rotor stages of the low-pressure turbine. 7.The turbomachine engine of claim 6, wherein the gearbox oil flow rate atthe inlet of the gearbox is within a range of 16-51 gallons per minuteat the max takeoff condition.
 8. The turbomachine engine of claim 7,wherein the diameter of the fan blades is 80-120 inches.
 9. Theturbomachine engine of claim 8, wherein the net thrust of theturbomachine engine is within a range of 25,000-80,000 pounds force atthe max takeoff condition.
 10. A turbomachine engine comprising: anacelle; a fan assembly disposed within the nacelle and including aplurality of fan blades arranged in a single blade row; a vane assemblydisposed within the nacelle and including a plurality of vanes arrangedin a single vane row and disposed aft of the plurality of fan blades; acore engine including a first compressor section, a second compressorsection, a first turbine section, and a second turbine section; a firstshaft coupling the first turbine section to the first compressorsection; a second shaft coupling the second turbine section to thesecond compressor section; a gearbox including an input and an output,wherein the input of the gearbox is coupled to the first shaft andcomprises a first rotational speed, wherein the output of the gearbox iscoupled to the fan assembly and has a second rotational speed, which isless than the first rotational speed, and wherein a gear ratio of thegearbox is within a range of 3.2-4.0; and an overall engine efficiencyrating greater than or equal to 0.35GR^(1.5) and less than or equal to0.7GR^(1.5), wherein the overall engine efficiency rating equals${{Q\left( \frac{D^{1.56}}{T} \right)}^{1.53}N^{2}},$ wherein Q is agearbox oil flow rate at an inlet of the gearbox measured in gallons perminute at a max takeoff condition, wherein D is a diameter of the fanblades measured in inches, wherein T is a net thrust of the turbomachineengine measured in pounds force at the max takeoff condition, andwherein N is a number of rotating blade stages of the first turbinesection and equals 3 or
 4. 11. The turbomachine engine of claim 10,wherein the gearbox is an epicyclic gearbox comprising, and wherein thegear ratio of the gearbox is within a range of 3.4-4.0.
 12. Theturbomachine engine of claim 10, wherein the overall engine efficiencyrating is within a range of 2.0-3.0.
 13. The turbomachine engine ofclaim 10, wherein the overall engine efficiency rating is within a rangeof 3.0-4.0.
 14. The turbomachine engine of claim 10, wherein the overallengine efficiency rating is within a range of 4.0-5.6.
 15. Theturbomachine engine of claim 10, wherein the second compressor sectioncomprises 8-11 stages.
 16. The turbomachine engine of claim 10, whereinthe second compressor section comprises 9-10 stages.
 17. Theturbomachine engine of claim 10, wherein the fan assembly comprises12-18 fan blades, wherein the first compressor section comprises 1-5stages, wherein the second compressor section comprises 8-10 stages,wherein the second turbine section comprises 2 stages.
 18. Theturbomachine engine of claim 10, wherein the fan assembly comprises16-18 fan blades, wherein the first compressor section comprises 3-5stages, wherein the second compressor section comprises 8-9 stages,wherein the second turbine section comprises 2 stages.
 19. Theturbomachine engine of claim 10, wherein the fan assembly comprises16-18 fan blades, wherein the first compressor section comprises 2-3stages, wherein the second compressor section comprises 9-10 stages,wherein the second turbine section comprises 2 stages.
 20. Theturbomachine engine of claim 10, wherein the fan assembly comprises12-20 fan blades, wherein the first compressor section comprises 1-5stages, wherein the second compressor section comprises 8-11 stages,wherein the second turbine section comprises 1-2 stages.